Turbine component cooling hole within a microsurface feature that protects adjoining thermal barrier coating

ABSTRACT

Cooling holes in a turbine component, such as a blade, vane or combustor transition, are formed in and surrounded by a micro surface feature (MSF) that protects the adjoining thermal barrier coating (TBC) from delamination or crack propagation during the hole formation or during engine operation. The MSF effectively functions as a circumferential sleeve around the cooling hole margin so that relatively more friable TBC material that would otherwise define the cooling hole margin is not directly exposed to coolant fluid exhausting the hole, foreign object damage (FOD) or contact with cooling hole formation tooling when fabricating the hole through the TBC layer. The MSF is formed as a projection from the component substrate or during subsequent application of a metallic bond coat (BC) layer.

PRIORITY CLAIM AND CROSS-REFERENCE TO RELATED APPLICATIONS

This application is the U.S. National stage of the InternationalApplication No. PCT/US2015/016288, filed Feb. 18, 2015, which is hereinincorporated by reference in its entirety.

The International Application No. PCT/US2015/016288 claims priorityunder the following United States patent applications, the entirecontents of each of which is incorporated by reference herein:

“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE HAVING A FRANGIBLEOR PIXELATED NIB SURFACE”, filed Feb. 25, 2014, and assigned Ser. No.14/188,941; and

“TURBINE ABRADABLE LAYER WITH PROGRESSIVE WEAR ZONE MULTI LEVEL RIDGEARRAYS”, filed Feb. 25, 2014, and assigned Ser. No. 14/188,958.

A concurrently filed International Patent Application entitled “TURBINEABRADABLE LAYER WITH AIRFLOW DIRECTING PIXELATED SURFACE FEATUREPATTERNS”, docket number 2013P20413WO, and assigned serial number(unknown) is identified as a related application and is incorporated byreference herein.

TECHNICAL FIELD

The invention relates to combustion or steam turbine engines havingthermal barrier coating (TBC) layers and cooling holes on its componentsurfaces that are exposed to heated working fluids, such as combustiongasses or high pressure steam, including individual subcomponents thatincorporate such TBC layers, such as blades, vanes or combustortransitions. More specifically the invention relates to protection ofthe TBC layer adjoining a component cooling hole margin by forming thehole in a micro surface feature (MSF), which circumscribes the hole. TheMSF effectively forms a metallic boundary, which separates the adjoiningTBC layer from the hole margin.

BACKGROUND OF THE INVENTION

Known turbine engines, including gas turbine engines and steam turbineengines, incorporate shaft-mounted turbine blades circumferentiallycircumscribed by a turbine casing or housing. Hot gasses flowing pastthe turbine blades cause blade rotation that converts thermal energywithin the hot gasses to mechanical work, which is available forpowering rotating machinery, such as an electrical generator. Referringto FIGS. 1-6, known turbine engines, such as the gas turbine engine 80include a multi stage compressor section 82, a combustor section 84, amulti stage turbine section 86 and an exhaust system 88. Atmosphericpressure intake air is drawn into the compressor section 82 generally inthe direction of the flow arrows F along the axial length of the turbineengine 80. The intake air is progressively pressurized in the compressorsection 82 by rows rotating compressor blades and directed by matingcompressor vanes to the combustor section 84, where it is mixed withfuel and ignited. The ignited fuel/air mixture, now under greaterpressure and velocity than the original intake air, is directed by acombustor transition 85 to the sequential rows R₁, R₂, etc., in theturbine section 86. The engine's rotor and shaft 90 has a plurality ofrows of airfoil cross sectional shaped turbine blades 92 terminating indistal blade tips 94 in the compressor 82 and turbine 86 sections. Forconvenience and brevity further discussion of turbine blades andabradable layers in the engine will focus on the turbine section 86embodiments and applications, though similar constructions areapplicable for the compressor section 82. Each blade 92 has a concaveprofile high pressure side 96 and a convex low pressure side 98. Thehigh velocity and pressure combustion gas, flowing in the combustionflow direction F imparts rotational motion on the blades 92, spinningthe rotor. As is well known, some of the mechanical power imparted onthe rotor shaft is available for performing useful work. The combustiongasses are constrained radially distal the rotor by turbine casing 100and proximal the rotor by air seals 102. Referring to the Row 1 sectionshown in FIG. 2, respective upstream vanes 104 and downstream vanes 106direct upstream combustion gas generally parallel to the incident angleof the leading edge of turbine blade 92 and redirect downstreamcombustion gas exiting the trailing edge of the blade. The enginecomponent surfaces that are in contact with working fluid combustiongasses often incorporate a thermal barrier coating (TBC) layer that isdirectly applied to the component substrate or over an intermediatemetallic bond coat that was previously applied over the substrate.

Turbine blades and vanes, especially in the engine turbine section 86,as well as combustor transitions in the combustor section 84 oftenincorporate cooling holes that are in communication with the workingfluid combustion gasses, in addition to a TBC layer. The cooling holesare formed through the TBC layer, exposing TBC material around theperiphery of the cooling hole margins. The TBC material is relativelymore friable and brittle than the component underlying metallicsubstrate BC layer. Thus TBC material adjoining cooling holes issusceptible to spallation or crack propagation during formation of thecooling hole through the TBC layer or during subsequent engineoperation.

By way of general background not directly relevant to cooling holesformed within TBC layered surfaces of turbine engine components, butnonetheless relevant to engine operation, the exemplary turbine engine80 turbine casing 100 proximal the blade tips 94 is lined with aplurality of sector shaped abradable components 110, each having asupport surface 112 retained within and coupled to the casing and anabradable substrate 120 that is in opposed, spaced relationship with theblade tip by a blade tip gap G. The abradable substrate is oftenconstructed of a metallic/ceramic material that has high thermal andthermal erosion resistance and that maintains structural integrity athigh combustion temperatures. As the abradable surface 120 metallicceramic materials is often more abrasive than the turbine blade tip 94material a blade tip gap G is maintained to avoid contact between thetwo opposed components that might at best cause premature blade tip wearand in worse case circumstances might cause engine damage. Some knownabradable components 110 are constructed with a monolithicmetallic/ceramic abradable substrate 120. Other known abradablecomponents 110 are constructed with a composite matrix composite (CMC)structure, comprising a ceramic support surface 112 to which is bonded afriable graded insulation (FGI) ceramic strata of multiple layers ofclosely-packed hollow ceramic spherical particles, surrounded by smallerparticle ceramic filler, as described in U.S. Pat. No. 6,641,907.Spherical particles having different properties are layered in thesubstrate 120, with generally more easily abradable spheres forming theupper layer to reduce blade tip 94 wear. Another CMC structure isdescribed in U.S. Patent Publication No. 2008/0274336, wherein thesurface includes a cut grooved pattern between the hollow ceramicspheres. The grooves are intended to reduce the abradable surfacematerial cross sectional area to reduce potential blade tip 94 wear, ifthey contact the abradable surface. Other commonly known abradablecomponents 110 are constructed with a metallic base layer supportsurface 112 to which is applied a thermally sprayed ceramic/metalliclayer that forms the abradable substrate layer 120. As will be describedin greater detail the thermally sprayed metallic layer may includegrooves, depressions or ridges to reduce abradable surface materialcross section for potential blade tip 94 wear reduction.

In addition to the desire to prevent blade tip 94 premature wear orcontact with the abradable substrate 120, as shown in FIG. 3, for idealairflow and power efficiency each respective blade tip 94 desirably hasa uniform blade tip gap G relative to the abradable component 110 thatis as small as possible (ideally zero clearance) to minimize blade tipairflow leakage L between the high pressure blade side 96 and the lowpressure blade side 98 as well as axially in the combustion flowdirection F. However, manufacturing and operational tradeoffs requireblade tip gaps G greater than zero. Such tradeoffs include tolerancestacking of interacting components, so that a blade constructed on thehigher end of acceptable radial length tolerance and an abradablecomponent abradable substrate 120 constructed on the lower end ofacceptable radial tolerance do not impact each other excessively duringoperation. Similarly, small mechanical alignment variances during engineassembly can cause local variations in the blade tip gap. For example ina turbine engine of many meters axial length, having a turbine casingabradable substrate 120 inner diameter of multiple meters, very smallmechanical alignment variances can impart local blade tip gap Gvariances of a few millimeters.

During turbine engine 80 operation the turbine engine casing 100 mayexperience out of round (e.g., egg shaped) thermal distortion as shownin FIGS. 4 and 6. Casing 100 thermal distortion potential increasesbetween operational cycles of the turbine engine 80 as the engine isfired up to generate power and subsequently cooled for servicing afterthousands of hours of power generation. Commonly, as shown in FIG. 6,greater casing 100 and abradable component 110 distortion tends to occurat the uppermost 122 and lowermost 126 casing circumferential positions(i.e., 6:00 and 12:00 positions) compared to the lateral right 124 andleft 128 circumferential positions (i.e., 3:00 and 9:00). If, forexample as shown in FIG. 4 casing distortion at the 6:00 position causesblade tip contact with the abradable substrate 120 one or more of theblade tips may be worn during operation, increasing the blade tip gaplocally in various other less deformed circumferential portions of theturbine casing 100 from the ideal gap G to a larger gap G_(W) as shownin FIG. 5. The excessive blade gap G_(w) distortion increases blade tipleakage L, diverting hot combustion gas away from the turbine blade 92airfoil, reducing the turbine engine's efficiency.

In the past flat abradable surface substrates 120 were utilized and theblade tip gap G specification conservatively chosen to provide at leasta minimal overall clearance to prevent blade tip 94 and abradablesurface substrate contact within a wide range of turbine componentmanufacturing tolerance stacking, assembly alignment variances, andthermal distortion. Thus, a relatively wide conservative gap Gspecification chosen to avoid tip/substrate contact sacrificed engineefficiency. Commercial desire to enhance engine efficiency for fuelconservation has driven smaller blade tip gap G specifications:preferably no more than 2 millimeters and desirably approaching 1millimeter.

Past abradable designs have incorporated rows of radially repeatingcontinuous ribs spanning the axial swept area of the blade tip with gapsbetween successive ribs, in order to reduce the potential surfacecontact area between the abradable ribs and the turbine blade tips. Theprojecting ribs were configured to control or inhibit hot gas flowacross the blade tip from the pressure to suction side of the tip. Forexample, in order to reduce likelihood of blade tip/substrate contact,abradable components comprising metallic base layer supports withthermally sprayed metallic/ceramic abradable surfaces have beenconstructed with three dimensional planform profiles, such as shown inFIGS. 7-11. The exemplary known abradable surface component 130 of FIGS.7 and 10 has a metallic base layer support 131 for coupling to a turbinecasing 100, upon which a thermally sprayed metallic/ceramic layer hasbeen deposited and formed into three-dimensional ridge and grooveprofiles by known deposition or ablative material working methods.Specifically in these cited figures a plurality of ridges 132respectively have a common height H_(R) distal ridge tip surface 134that defines the blade tip gap G between the blade tip 94 and it. Eachridge also has side walls 135 and 136 that extend from the substratesurface 137 and define grooves 138 between successive ridge opposed sidewalls. The ridges 132 are arrayed with parallel spacing S_(R) betweensuccessive ridge center lines and define groove widths W_(G). Due to theabradable component surface symmetry, groove depths D_(G) correspond tothe ridge heights H_(R). Compared to a solid smooth surface abradable,the ridges 132 have smaller cross section and more limited abrasioncontact in the event that the blade tip gap G becomes so small as toallow blade tip 94 to contact one or more tips 134. However therelatively tall and widely spaced ridges 132 allow blade leakage L intothe grooves 138 between ridges, as compared to the prior continuous flatabradable surfaces. In an effort to reduce blade tip leakage L, theridges 132 and grooves 138 were oriented horizontally in the directionof combustion flow F (not shown) or diagonally across the width of theabradable surface 137, as shown in FIG. 7, so that they would tend toinhibit the leakage. Other known abradable components 140, shown in FIG.8, have arrayed grooves 148 in crisscross patterns, forming diamondshaped ridge planforms 142 with flat, equal height ridge tips 144.Additional known abradable components have employed triangular roundedor flat tipped triangular ridges 152 shown in FIGS. 9 and 11. In theabradable component 150 of FIGS. 9 and 11, each ridge 152 hassymmetrical side walls 155, 156 that terminate in a flat ridge tip 154.All ridge tips 154 have a common height H_(R) and project from thesubstrate surface 157. Grooves 158 are curved and have a similarplanform profile as the blade tip 94 camber line. Curved grooves 158generally are more difficult to form than linear grooves 138 or 148 ofthe abradable components shown in FIGS. 7 and 8.

Past abradable component designs have required stark compromises betweenblade tips wear resulting from contact between the blade tip and theabradable surface and blade tip leakage that reduces turbine engineoperational efficiency. Optimizing engine operational efficiencyrequired reduced blade tip gaps and smooth, consistently flat abradablesurface topology to hinder air leakage through the blade tip gap,improving initial engine performance and energy conservation. Aspreviously noted, any gap between the tip of a rotating blade and thesurface to which it seals will result in a loss of turbine efficiencydue to the depressurization of hot gas flowing over the tip of the bladerather than through the turbine. Abradable systems have finite servicelives that are primarily attributable to either increased hardness ofthe abradable through gradual sintering by rubbing against the blade tipor loss of the coating through spallation. It is desirable to balancesmall blade tip/abradable surface gap and low erosion of those opposedsurfaces for longer turbine service life between service outages.

In another drive for increased gas turbine operational efficiency andflexibility so-called “fast start” mode engines were being constructedthat required faster full power ramp up (order of 40-50 Mw/minute).Aggressive ramp-up rates exacerbated potential higher incursion of bladetips into ring segment abradable coating, resulting from quicker thermaland mechanical growth and higher distortion and greater mismatch ingrowth rates between rotating and stationary components. This in turnrequired greater turbine tip clearance in the “fast start” mode engines,to avoid premature blade tip wear, than the blade tip clearance requiredfor engines that are configured only for “standard” starting cycles.Thus as a design choice one needed to balance the benefits of quickerstartup/lower operational efficiency larger blade tip gaps or standardstartup/higher operational efficiency smaller blade tip gaps.Traditionally standard or fast start engines required differentconstruction to accommodate the different needed blade tip gapparameters of both designs. Whether in standard or fast startconfiguration, decreasing blade tip gap for engine efficiencyoptimization ultimately risked premature blade tip wear, opening theblade tip gap and ultimately decreasing longer term engine performanceefficiency during the engine operational cycle. The aforementionedceramic matrix composite (CMC) abradable component designs sought tomaintain airflow control benefits and small blade tip gaps of flatsurface profile abradable surfaces by using a softer top abradable layerto mitigate blade tip wear. The abradable components of the U.S. PatentPublication No. 2008/0274336 also sought to reduce blade tip wear byincorporating grooves between the upper layer hollow ceramic spheres.However groove dimensions were inherently limited by the packing spacingand diameter of the spheres in order to prevent sphere breakage. Addinguniform height abradable surface ridges to thermally sprayed substrateprofiles as a compromise solution to reduce blade tip gap while reducingpotential rubbing contact surface area between the ridge tips and bladetips reduced likelihood of premature blade tip wear/increasing blade tipgap but at the cost of increased blade tip leakage into grooves betweenridges. As noted above, attempts have been made to reduce blade tipleakage flow by changing planform orientation of the ridge arrays toattempt to block or otherwise control leakage airflow into the grooves.

SUMMARY OF THE INVENTION

In various embodiments of the invention that are described herein, steamor combustion turbine engine components, such as blades, vanes orcombustor transitions, are constructed with cooling holes that areformed in and surrounded by a micro surface feature (MSF). The MSFprotects the adjoining thermal barrier coating (TBC) from delaminationor crack propagation during the hole formation process or during engineoperation. The MSF effectively functions as a circumferential sleevearound the cooling hole margin so that relatively more friable TBCmaterial that would otherwise define the cooling hole margin is notdirectly exposed to coolant fluid exhausting the hole, foreign objectdamage (FOD) or contact with cooling hole formation tooling whenfabricating the hole through the TBC layer. The MSF is formed as aprojection from the component substrate or during subsequent applicationof a metallic bond coat (BC) layer.

More particularly, exemplary embodiments of the invention feature aturbine component that is adapted for incorporation within a turbineengine, having an outer surface for exposure to heated working fluidthat drives the engine (such as combustion gas within a combustionturbine engine). The component includes a metallic substrate having asubstrate surface. A micro surface feature (MSF) projects from thesubstrate surface, having an MSF sidewall and an MSF upper surfaceforming part of the turbine component outer surface that caps the MSFsidewall. A cooling hole is formed within and is circumscribed by theMSF upper surface, with the hole extending within the substrate. Athermally sprayed or vapor deposited or solution/suspension plasmasprayed thermal barrier coat (TBC) is applied over the substrate andabutting the MSF sidewall, forming part of the component outer surface,for exposure to heated working fluid. In some embodiments of theinvention the cooling hole or the MSF sidewall or both have central axesthat are skewed relative to the substrate surface. In other embodimentsthe MSF sidewall has an undercut outer surface for mechanicallyanchoring the TBC layer to the MSF. In some embodiments the MSF isformed in the metallic substrate, while in other embodiments the MSF isformed in or covered by a bond coat (BC) that is interposed between thesubstrate and the TBC layer. The component may comprise a plurality ofthe MSFs with cooling holes therein that are arrayed about thesubstrate.

Other embodiments of the invention are directed to a turbine engine thatincludes a turbine housing; a rotor having blades rotatively mounted inthe turbine housing, a rotor having blades rotatively mounted in theturbine housing; and turbine vanes mounted in the turbine housing atleast upstream of the blades. At least one turbine component has anouter surface for exposure to heated working fluid that drives theengine (such as combustion gas within a combustion turbine engine). Thecomponent includes a metallic substrate having a substrate surface. Amicro surface feature (MSF) projects from the substrate surface, havingan MSF sidewall and an MSF upper surface forming part of the turbinecomponent outer surface that caps the MSF sidewall. A cooling hole isformed within and is circumscribed by the MSF upper surface, with thehole extending within the substrate. A thermally sprayed or vapordeposited or solution/suspension plasma sprayed thermal barrier coat(TBC) is applied over the substrate and abutting the MSF sidewall,forming part of the component outer surface, for exposure to heatedworking fluid. In some embodiments of the invention the cooling hole orthe MSF sidewall or both have central axes that are skewed relative tothe substrate surface. In other embodiments the MSF sidewall has anundercut outer surface for mechanically anchoring the TBC layer to theMSF. In some embodiments the MSF is formed in the metallic substrate,while in other embodiments the MSF is formed in or covered by a bondcoat (BC) that is interposed between the substrate and the TBC layer.The component may comprise a plurality of the MSFs with cooling holestherein that are arrayed about the substrate.

Yet other embodiments of the invention are directed to a method formaking a turbine component that is adapted for incorporation within aturbine engine, having an outer surface for exposure to heated workingfluid that drives the engine and cooling holes formed through the outersurface. A metallic substrate having a substrate surface is provided. Amicro surface feature (MSF) is formed on and projects from the substratesurface. The MSF has an MSF sidewall and an MSF upper surface. The MSFupper surface forms part of the turbine component outer surface, cappingthe MSF sidewall. A thermally sprayed or vapor deposited orsolution/suspension plasma deposited thermal barrier coat (TBC) layer isapplied over the substrate surface, abutting the MSF sidewall. The TBClayer forms part of the component outer surface, for exposure to engineheated working fluid. A cooling hole is formed within and iscircumscribed by the MSF upper surface.

The respective features of the invention may be applied jointly orseverally in any combination or sub-combination by those skilled in theart.

BRIEF DESCRIPTION OF THE DRAWINGS

The teachings of the invention can be readily understood by consideringthe following detailed description in conjunction with the accompanyingdrawings, in which:

FIG. 1 is a partial axial cross sectional view of an exemplary known gasturbine engine;

FIG. 2 is a detailed cross sectional elevational view of Row 1 turbineblade and vanes showing blade tip gap G between a blade tip andabradable component of the turbine engine of FIG. 1;

FIG. 3 is a radial cross sectional schematic view of a known turbineengine, with ideal uniform blade tip gap G between all blades and allcircumferential orientations about the engine abradable surface;

FIG. 4 is a radial cross sectional schematic view of an out of roundknown turbine engine showing blade tip and abradable surface contact atthe 12:00 uppermost and 6:00 lowermost circumferential positions;

FIG. 5 is a radial cross sectional schematic view of a known turbineengine that has been in operational service with an excessive blade tipgap G_(w) that is greater than the original design specification bladetip gap G;

FIG. 6 is a radial cross sectional schematic view of a known turbineengine, highlighting circumferential zones that are more likely tocreate blade tip wear and zones that are less likely to create blade tipwear;

FIGS. 7-9 are plan or plan form views of known ridge and groove patternsfor turbine engine abradable surfaces;

FIGS. 10 and 11 are cross sectional elevational views of known ridge andgroove patterns for turbine engine abradable surfaces taken alongsections C-C of FIGS. 7 and 9, respectively;

FIGS. 12-17 are plan or plan form views of “hockey stick” configurationridge and groove patterns of turbine engine abradable surfaces, inaccordance with exemplary embodiments that are described in greaterdetail herein, with schematic overlays of turbine blades;

FIGS. 18 and 19 are plan or plan form views of another “hockey stick”configuration ridge and groove pattern for a turbine engine abradablesurface that includes vertically oriented ridge or rib arrays alignedwith a turbine blade rotational direction, in accordance with anotherexemplary embodiment, and a schematic overlay of a turbine blade;

FIG. 20 is a comparison graph of simulated blade tip leakage mass fluxfrom leading to trailing edge for a respective exemplary continuousgroove hockey stick abradable surface profile of the type shown in FIGS.12-17 and a split groove with interrupting vertical ridges hockey stickabradable surface profile of the type shown in FIGS. 18 and 19;

FIG. 21 is a plan or plan form view of another “hockey stick”configuration ridge and groove pattern for an abradable surface, havingintersecting ridges and grooves, in accordance with another exemplaryembodiment, and a schematic overlay of a turbine blade;

FIG. 22 is a plan or plan form view of another “hockey stick”configuration ridge and groove pattern for an abradable surface, similarto that of FIGS. 18 and 19, which includes vertically oriented ridgearrays that are laterally staggered across the abradable surface in theturbine engine's axial flow direction, in accordance with anotherexemplary embodiment;

FIG. 23 is a plan or plan form view of a “zig-zag” configuration ridgeand groove pattern for an abradable surface, which includes horizontallyoriented ridge and groove arrays across the abradable surface in theturbine engine's axial flow direction, in accordance with anotherexemplary embodiment;

FIG. 24 is a plan or plan form view of a “zig-zag” configuration ridgeand groove pattern for an abradable surface, which includes diagonallyoriented ridge and groove arrays across the abradable surface, inaccordance with another exemplary embodiment;

FIG. 25 is a plan or plan form view of a “zig-zag” configuration ridgeand groove pattern for an abradable surface, which includes Vee shapedridge and groove arrays across the abradable surface, in accordance withanother exemplary embodiment;

FIGS. 26-29 are plan or plan form views of nested loop configurationridge and groove patterns of turbine engine abradable surfaces, inaccordance with exemplary embodiments, with schematic overlays ofturbine blades;

FIGS. 30-33 are plan or plan form views of maze or spiral configurationridge and groove patterns of turbine engine abradable surfaces, inaccordance with exemplary embodiments, with schematic overlays ofturbine blades;

FIGS. 34 and 35 are plan or plan form views of a compound angle withcurved rib transitional section configuration ridge and groove patternfor a turbine engine abradable, in accordance with another exemplaryembodiment, and a schematic overlay of a turbine blade;

FIG. 36 is a comparison graph of simulated blade tip leakage mass fluxfrom leading to trailing edge for a respective exemplary compound anglewith curved rib transitional section configuration ridge and groovepattern abradable surface of the type of FIGS. 34 and 35, an exemplaryknown diagonal ridge and groove pattern of the type shown in FIG. 7, anda known axially aligned ridge and groove pattern abradable surfaceabradable surface profile;

FIG. 37 is a plan or plan form view of a multi height or elevation ridgeprofile configuration and corresponding groove pattern for an abradablesurface, suitable for use in either standard or “fast start” enginemodes, in accordance with an exemplary embodiment;

FIG. 38 is a cross sectional view of the abradable surface embodiment ofFIG. 37 taken along C-C thereof;

FIG. 39 is a schematic elevational cross sectional view of a movingblade tip and abradable surface embodiment of FIGS. 37 and 38, showingblade tip leakage L and blade tip boundary layer flow in accordance withembodiments described herein;

FIGS. 40 and 41 are schematic elevational cross sectional views similarto FIG. 39, showing blade tip gap G, groove and ridge multi height orelevational dimensions in accordance with embodiments described herein;

FIG. 42 is an elevational cross sectional view of a known abradablesurface ridge and groove profile similar to FIG. 11;

FIG. 43 is an elevational cross sectional view of a multi height orelevation stepped profile ridge configuration and corresponding groovepattern for an abradable surface, in accordance with an exemplaryembodiment;

FIG. 44 is an elevational cross sectional view of another embodiment ofa multi height or elevation stepped profile ridge configuration andcorresponding groove pattern for an abradable surface of the invention;

FIG. 45 is an elevational cross sectional view of a multi depth grooveprofile configuration and corresponding ridge pattern for an abradablesurface, in accordance with an embodiment described herein;

FIG. 46 is an elevational cross sectional view of an asymmetric profileridge configuration and corresponding groove pattern for an abradablesurface, in accordance with an embodiment described herein;

FIG. 47 a perspective view of an asymmetric profile ridge configurationand multi depth parallel groove profile pattern for an abradablesurface, in accordance with an embodiment described herein;

FIG. 48 is a perspective view of an asymmetric profile ridgeconfiguration and multi depth intersecting groove profile pattern for anabradable surface, wherein upper grooves are tipped longitudinallyrelative to the ridge tip, in accordance with an embodiment describedherein;

FIG. 49 is a perspective view of another embodiment of an asymmetricprofile ridge configuration and multi depth intersecting groove profilepattern for an abradable surface, wherein upper grooves are normal toand skewed longitudinally relative to the ridge tip;

FIG. 50 is an elevational cross sectional view of cross sectional viewof a multi depth, parallel groove profile configuration in a symmetricprofile ridge for an abradable surface, in accordance with anotherembodiment;

FIGS. 51 and 52 are respective elevational cross sectional views ofmulti depth, parallel groove profile configurations in a symmetricprofile ridge for an abradable surface, wherein an upper groove istilted laterally relative to the ridge tip, in accordance with anembodiment described herein;

FIG. 53 is a perspective view of an abradable surface, in accordancewith embodiment, having asymmetric, non-parallel wall ridges and multidepth grooves;

FIGS. 54-56 are respective elevational cross sectional views of multidepth, parallel groove profile configurations in a trapezoidal profileridge for an abradable surface, wherein an upper groove is normal to ortilted laterally relative to the ridge tip, in accordance withalternative embodiments described herein;

FIG. 57 is a is a plan or plan form view of a multi-level intersectinggroove pattern for an abradable surface in accordance with an embodimentdescribed herein;

FIG. 58 is a perspective view of a stepped profile abradable surfaceridge, wherein the upper level ridge has an array of pixelatedupstanding nibs projecting from the lower ridge plateau, in accordancewith an embodiment described herein;

FIG. 59 is an elevational view of a row of pixelated upstanding nibsprojecting from the lower ridge plateau, taken along C-C of FIG. 58;

FIG. 60 is an alternate embodiment of the upstanding nibs of FIG. 59,wherein the nib portion proximal the nib tips are constructed of a layerof material having different physical properties than the material belowthe layer, in accordance with an embodiment described herein;

FIG. 61 is a schematic elevational view of the pixelated upper nibembodiment of FIG. 58, wherein the turbine blade tip deflects the nibsduring blade rotation;

FIG. 62 is a schematic elevational view of the pixelated upper nibembodiment of FIG. 58, wherein the turbine blade tip shears off all or apart of upstanding nibs during blade rotation, leaving the lower ridgeand its plateau intact and spaced radially from the blade tip by a bladetip gap;

FIG. 63 is a schematic elevational view of the pixelated upper nibembodiment of FIG. 58, wherein the turbine blade tip has sheared off allof the upstanding nibs during blade rotation and is abrading the plateausurface of the lower ridge portion;

FIG. 64 is a plan or planform view of peeled layers of an abradablecomponent with a curved elongated pixelated major planform pattern(PMPP) of a plurality of micro surface features (MSF), in accordancewith an exemplary embodiment described herein;

FIG. 65 is a plan or planform view of peeled layers of an abradablecomponent with a diagonal elongated pixelated major planform pattern(PMPP) of a plurality of micro surface features (MSF), in accordancewith another exemplary embodiment described herein;

FIG. 66 is a plan or planform view showing peeled layers of an abradablecomponent with a “hockey-stick” elongated pixelated major planformpattern (PMPP) of a plurality of micro surface features (MSF), inaccordance with another exemplary embodiment;

FIG. 67 is a fragmented plan or planform view showing an abradablecomponent surface with a herringbone pixelated major planform pattern(PMPP) of a plurality of chevron-shaped micro surface features (MSF), inaccordance with an exemplary embodiment;

FIG. 68 is a detailed perspective view of a chevron-shaped micro surfacefeature (MSF) of FIG. 67;

FIG. 69 is a fragmented plan or planform view showing an abradablecomponent surface with a herringbone pixelated major planform pattern(PMPP) of a plurality of an alternative embodiment chevron-shaped microsurface features (MSF), which comprise two linear elements converging atan apex that are separated by a gap at the apex;

FIG. 70 is a detailed perspective view of the alternative embodimentchevron-shaped micro surface feature (MSF) of FIG. 69;

FIG. 71 is a fragmented plan or planform view showing an abradablecomponent surface with a pixelated major planform pattern (PMPP) of aplurality of curved- or annular sector-shaped micro surface features(MSF), in accordance with an exemplary embodiment;

FIG. 72 is a detailed perspective view of an annular sector-shaped microsurface feature (MSF) of FIG. 71;

FIG. 73 is a fragmented plan or planform view showing an abradablecomponent surface with a pixelated major planform pattern (PMPP) ofcomposite annular sector-shaped and rectangular or linear micro surfacefeatures (MSF), in accordance with an exemplary embodiment;

FIG. 74 is a detailed perspective view of the composite annularsector-shaped and linear micro surface features (MSF) of FIG. 73;

FIG. 75 is a fragmented plan or planform view showing an abradablecomponent surface with a diamond pixelated major planform pattern (PMPP)of linear micro surface features (MSF), in accordance with an exemplaryembodiment;

FIG. 76 is a fragmented plan or planform view showing an abradablecomponent surface with a undulating pattern pixelated major planform(PMPP) of curved micro surface features (MSF), in accordance with anexemplary embodiment;

FIG. 77 is a fragmented plan or planform view showing an abradablecomponent surface with a pixelated major planform pattern (PMPP) ofdiscontinuous curved micro surface features (MSF), in accordance with anexemplary embodiment;

FIG. 78 is a fragmented plan or planform view showing an abradablecomponent surface with a zig-zag undulating pixelated major planformpattern (PMPP) of first height and higher second height micro surfacefeatures (MSF), in accordance with an exemplary embodiment;

FIG. 79 is a cross sectional view of the abradable component of FIG. 78;

FIG. 80 is a fragmented plan or planform view showing an abradablecomponent surface with a zig-zag undulating pixelated major planformpattern (PMPP) of first height and higher second height micro surfacefeatures (MSF), in accordance with another exemplary embodiment;

FIG. 81 is a cross sectional view of the abradable component of FIG. 80;

FIG. 82 is a cross sectional view of an abradable component with microsurface features (MSF) formed in a metallic bond coat that is appliedover a support substrate, in accordance with an exemplary embodiment;

FIG. 83 is a cross sectional view of an abradable component with microsurface features (MSF) formed in a support substrate, in accordance withanother exemplary embodiment;

FIG. 84 is a detailed cross sectional elevational view, similar to thatof FIG. 2, of a turbine engine with Row 1 turbine blade and Rows 1 and 2vanes incorporating one or more exemplary cooling hole micro surfacefeature (MSF) embodiments of the invention;

FIG. 85 is an exterior plan view of a turbine engine component coolinghole within an MSF, oriented on an outer surface of the component thatis exposed to hot working fluid gas in the engine, in accordance with anexemplary embodiment of the invention;

FIG. 86 is a cross sectional view of the cooling hole within an MSF ofFIG. 85, wherein the MSF is formed directly on the component substrate;

FIGS. 87-89 are cross sectional views of alternative embodiment coolingholes within MSFs;

FIGS. 90-94 are cross sectional views of exemplary method steps formaking a cooling hole within an MSF, in accordance with embodiments ofthe invention, wherein as in FIG. 86, the MSF is formed directly on thesubstrate;

FIG. 95 is a cross sectional view of the cooling hole within an MSF,wherein the MSF is formed in a bond coat (BC) layer applied over thecomponent substrate; and

FIGS. 96-98 are cross sectional views of alternative exemplary methodsteps for making a cooling hole within an MSF, in accordance withembodiments of the invention, wherein as in FIG. 95, the MSF is formedin a bond coat (BC) layer applied over the component substrate.

To facilitate understanding, identical reference numerals have beenused, where possible, to designate identical elements that are common tothe figures. The figures are not drawn to scale. The following commondesignators for dimensions, cross sections, fluid flow, turbine bladerotation, axial or radial orientation and fluid pressure have beenutilized throughout the various invention embodiments described herein:

A forward or upstream zone of an abradable surface;aft or downstream zone of an abradable surface;C-C abradable cross section;D_(G) abradable groove depth;F flow direction through turbine engine;G turbine blade tip to abradable surface gap;G_(W) worn turbine blade tip to abradable surface gap;H height of a micro surface feature (MSF);H_(R) abradable ridge height;L turbine blade tip leakage or length of a micro surface feature (MSF);P abradable surface plan view or planform;P_(P) turbine blade higher pressure side;P_(S) turbine blade lower pressure or suction side;R turbine blade rotational direction;R₁ Row 1 of the turbine engine turbine section;R₂ Row 2 of the turbine engine turbine section;S_(R) abradable ridge centerline spacing;W width of a micro surface feature (MSF);W_(G) abradable groove width;W_(R) abradable ridge width;α abradable groove planform angle relative to the turbine engine axialdimension;β abradable ridge sidewall angle relative to vertical or normal theabradable surface;γ abradable groove fore-aft tilt angle relative to abradable ridgeheight;Δ abradable groove skew angle relative to abradable ridge longitudinalaxis;ε abradable upper groove tilt angle relative to abradable surface and/orridge surface; andΦ abradable groove arcuate angle.

DESCRIPTION OF EMBODIMENTS

Embodiments of the invention described herein can be readily utilized inturbine engine components, including gas turbine engines, wherecomponent surfaces that are exposed to hot working fluid, such ascombustion gas, have thermal barrier coatings (TBCs) and cooling holesformed in the TBC layer. In exemplary embodiments described in greaterdetail herein, cooling holes in a turbine component, such as a blade,vane or combustor transition, are formed in and surrounded by a microsurface feature (MSF) “sleeve” that protects the adjoining thermalbarrier coating (TBC) from delamination or crack propagation during thehole formation or during engine operation. The MSF effectively functionsas a circumferential sleeve around the cooling hole margin so thatrelatively more friable TBC material that would otherwise define thecooling hole margin is not directly exposed to coolant fluid exhaustingthe hole, foreign object damage (FOD) or contact with cooling holeformation tooling when fabricating the hole through the TBC layer. TheMSF is formed as a projection from the component substrate or duringsubsequent application of a metallic bond coat (BC) layer. The microsurface features (MSFs) are formed by: (i) known thermal spray of moltenparticles to build up the surface feature or (ii) known additive layermanufacturing build-up application of the surface feature, such as by3-D printing, sintering, electron or laser beam deposition or (iii)known ablative removal of substrate material manufacturing processes,defining the feature by portions that were not removed.

Features of various embodiments of the invention that are describedherein can be combined to satisfy performance requirements of differentturbine applications, even though not every possible combination ofembodiments and features of the invention is specifically described indetail herein.

General Summary of Thermally Sprayed TBC Application in CombustionTurbine Engine Components

The turbine engine of FIG. 84 includes cooling holes 85A/99/105 withinturbine component outer surfaces that are constructed in accordance withexemplary embodiments of the present invention, which are not shown inthe turbine engine of FIGS. 1 and 2. For simplicity and ease ofcomprehension, identical reference numbers used for equivalentcomponents shown in the respective sets of figures. Referring to FIG.84, the combustion turbine engine 80 includes a multi stage compressorsection 82, a combustion section 84, a multi stage turbine section 86and an exhaust system 88. Atmospheric pressure intake air is drawn intothe compressor section 82 generally in the direction of the flow arrowsF along the axial length of the turbine engine 80. The intake air isprogressively pressurized in the compressor section 82 by rows rotatingcompressor blades and directed by mating compressor vanes to thecombustion section 84, where it is mixed with fuel and ignited. Theignited fuel/air mixture, now under greater pressure and velocity thanthe original intake air, is directed through a transition 85 to thesequential blade rows R₁, R₂, etc., in the turbine section 86. Theengine's rotor and shaft 90 has a plurality of rows of airfoil crosssectional shaped turbine blades 92 terminating in distal blade tips 94in the compressor 82 and turbine 86 sections. For convenience andbrevity further discussion of thermal barrier coat (TBC) layers on theengine components will focus on the turbine section 86 embodiments andapplications, though similar constructions are applicable for thecompressor 82 or combustion 84 sections, as well as for steam turbineengine components. In the engine's 80 turbine section 86, each turbineblade 92 has a concave profile high pressure side 96 and a convex lowpressure side 98. Cooling holes 99 that are formed in the blade 92facilitate passage of cooling fluid along the blade surface. The highvelocity and pressure combustion gas, flowing in the combustion flowdirection F imparts rotational motion on the blades 92, spinning therotor. As is well known, some of the mechanical power imparted on therotor shaft is available for performing useful work. The combustiongasses are constrained radially distal the rotor by turbine casing 100and proximal the rotor by air seals 102 comprising abradable surfaces.Referring to the Row 1 section shown in FIG. 2, respective upstreamvanes 104 and downstream vanes 106 respectively direct upstreamcombustion gas generally parallel to the incident angle of the leadingedge of turbine blade 92 and redirect downstream combustion gas exitingthe trailing edge of the blade for a desired entry angle into downstreamRow 2 turbine blades (not shown). Cooling holes 105 that are formed inthe vanes 104, 106 facilitate passage of cooling fluid along the vanesurface. It is noted that the cooling holes 85A, 99 and 105 shown inFIG. 84 are merely schematic representations, are enlarged for visualclarity and are not drawn to scale. A typical combustor transition 85,turbine blade 92 or vane 104, 106 has many more cooling holesdistributed about their respective outer surfaces that are of muchsmaller diameter relative to the respective transition, blade or vanetotal surface area that is exposed to the engine combustion gas.

As previously noted, turbine component surfaces that are exposed tocombustion gasses are often constructed with a thermal barrier coating(TBC) layer for insulation of their underlying substrates. Typical TBCcoated surfaces include the turbine blades 92, the vanes 104, 106 andrelated turbine vane carrier surfaces and combustion section transitions85. The TBC layer for blade 92, vane 104, 106 and transition 85 exposedsurfaces are often applied by thermal sprayed or vapor deposition orsolution/suspension plasma spray methods, with a total TBC layerthickness of 300-2000 microns (μm).

FIGS. 12-41 and 43-83 are exemplary turbine blade tip opposing abradablesurface planform and projection profile invention embodiments describedin the related patent applications for which priority is claimed herein.The abradable component cross sectional profiles shown in FIGS. 38-56and 58-63 that are formed in the thermally sprayed or vapor depositedabradable layer comprise composite multi height/depth ridge and groovepatterns that have distinct upper (zone I) and lower (zone II) wearzones. The abradable component cross sectional profiles shown in FIGS.64-83 comprise pixelated major planform patterns (PMPP) of discontinuousmicro surface features (MSF), over which is applied an abradable layer,so that the finished blade tip abradable layer 120 has aggregateplanform and cross sectional patterns of ridge and groove patternssimilar to those of the solid rib and groove constructions of FIGS.12-37 and 57.

With respect to the FIGS. 12-37 and 57 abradable surface patterns—againwith ridges and grooves projecting multiple thousands of microns abovethe underlying substrate surface compared to 2000 or less TBC layerthickness on blade, vane or transition component combustion gas exposedsurfaces—the lower wear zone II optimizes engine airflow and structuralcharacteristics while the upper wear zone I minimizes blade tip gap andwear by being more easily abradable than the lower zone. Variousembodiments of the abradable component afford easier abradability of theupper zone with upper sub ridges or nibs having smaller cross sectionalarea than the lower zone rib structure. In some embodiments the uppersub ridges or nibs are formed to bend or otherwise flex in the event ofminor blade tip contact and wear down and/or shear off in the event ofgreater blade tip contact. In other embodiments the upper zone I subridges or nibs are pixelated into arrays of upper wear zones so thatonly those nibs in localized contact with one or more blade tips areworn while others outside the localized wear zone remain intact. In theevent that the localized blade tip gap is further reduced, the bladetips wear away the zone II lower ridge portion at that location. Howeverthe relatively higher ridges outside that lower ridge portion localizedwear area maintain smaller blade tip gaps to preserve engine performanceefficiency.

With the progressive wear zones construction of some blade tip abradablewear surface 120 embodiments of the prior applications for whichpriority is claimed herein, blade tip gap G can be reduced frompreviously acceptable known dimensions. For example, if a knownacceptable blade gap G design specification is 1 mm the higher ridges inwear zone I can be increased in height so that the blade tip gap isreduced to 0.5 mm. The lower ridges that establish the boundary for wearzone II are set at a height so that their distal tip portions are spaced1 mm from the blade tip. In this manner a 50% tighter blade tip gap G isestablished for routine turbine operation, with acceptance of somepotential wear caused by blade contact with the upper ridges in zone I.Continued localized progressive blade wearing in zone II will only beinitiated if the blade tip encroaches into the lower zone, but in anyevent the blade tip gap G of 1 mm is no worse than known blade tip gapspecifications. In some exemplary embodiments the upper zone I height isapproximately ⅓ to ⅔ of the lower zone II height. If the blade tip gap Gbecomes reduced for any one or more blades due to turbine casing 100distortion, fast engine startup mode or other reason initial contactbetween the blade tip 94 and the abradable component 10 will occur atthe higher ridge tips forming Zone I. While still in zone I the bladetips 94 only rub the alternate staggered higher ridges. If the blade gapG progressively becomes smaller, the higher ridges will be abraded untilthey are worn all the way through zone I and start to contact the lowerridge tips in zone II. Once in Zone II the turbine blade tip 94 rubs allof the remaining ridges at the localized wear zone, but in otherlocalized portions of the turbine casing there may be no reduction inthe blade tip gap G and the upper ridges may be intact at their fullheight. Thus the alternating height rib construction of some of theabradable component 110 embodiments accommodates localized wear withinzones I and II, but preserve the blade tip gap G and the aerodynamiccontrol of blade tip leakage in those localized areas where there is noturbine casing 100 or blade 92 distortion.

Multi-height wear zone constructions in abradable components are alsobeneficial for so-called “fast start” mode engines that require fasterfull power ramp up (order of 40-50 Mw/minute). Aggressive ramp-up ratesexacerbate potential higher incursion of blade tips into ring segmentabradable coating 120, resulting from quicker thermal and mechanicalgrowth and higher distortion and greater mismatch in growth ratesbetween rotating and stationary components. When either standard or faststart or both engine operation modes are desired the taller ridges ZoneI form the primary layer of clearance, with the smallest blade tip gapG, providing the best energy efficiency clearance for machines thattypically utilize lower ramp rates or that do not perform warm starts.Generally the ridge height for the lower ridge tips in Zone II isbetween 25%-75% of the higher ridge tip height of those forming Zone I.

Turbine Blade Tip Abradable Component TBC Application

Insulative layers of greater thickness than 1000 microns are oftenapplied to sector shaped turbine blade tip abradable components 110(hereafter referred to generally as an “abradable component”) that linethe turbine engine 80 turbine casing 100 in opposed relationship withthe blade tips 94. The abradable components 110 having a support surface112 retained within and coupled to the casing and an insulativeabradable substrate 120 that is in opposed, spaced relationship with theblade tip by a blade tip gap G. The abradable substrate is oftenconstructed of a metallic/ceramic material, similar to the TBC coatingmaterials that are applied to blade 92, vane 104, 106 and transition 85combustion gas exposed surfaces. Those abradable substrate materialshave high thermal and thermal erosion resistance and maintain structuralintegrity at high combustion temperatures. Generally, it should beunderstood that some form of TBC layer is formed over the blade tipabradable component 110 bare underlying metallic support surfacesubstrate 112 for the latter's insulative protection plus the insulativesubstrate thickness that projects at additional height over the TBC.Thus it should be understood that abradable components 110 have afunctionally equivalent TBC layer to the TBC layer applied over theturbine transition 85, blade 92 and vane 102/104, The abradable surface120 function is analogous to a shoe sole or heel that protects theabradable component support surface substrate 112 from wear and providesan additional layer of thermal protection. Exemplary materials used forblade tip abradable surface ridges/grooves include pyrochlore, fullycubic or partially stabilized yttria stabilized zirconia. As theabradable surface 120 metallic ceramic materials is often more abrasivethan the turbine blade tip 94 material a blade tip gap G is maintainedto avoid contact between the two opposed components that might at bestcause premature blade tip wear and in worse case circumstances mightcause engine damage.

Blade tip abradable components 110 are often constructed with a metallicbase layer support surface 112, to which is applied a thermally sprayedceramic/metallic abradable substrate layer 120 of many thousands ofmicrons thickness, i.e., multiples of the typical transition 85 blade 92or vane 104/106 TBC layer thickness. As will be described in greaterdetail herein, the abradable layer of exemplary turbine blade tipopposing abradable surface planform and projection profile inventionembodiments described in the related patent applications for whichpriority is claimed herein include grooves, depressions or ridges in theabradable substrate layer 120 to reduce abradable surface material crosssection for potential blade tip 94 wear reduction and for directingcombustion airflow in the gap region G. Commercial desire to enhanceengine efficiency for fuel conservation has driven smaller blade tip gapG specifications: preferably no more than 2 millimeters and desirablyapproaching 1 millimeter (1000 μm).

Abradable Surface Planforms

Exemplary invention embodiment abradable surface ridge and grooveplanform patterns are shown in FIGS. 12-37 and 57. Unlike knownabradable planform patterns that are uniform across an entire abradablesurface, many of the present invention planform pattern embodiments arecomposite multi groove/ridge patterns that have distinct forwardupstream (zone A) and aft downstream patterns (zone B). Those combinedzone A and zone B ridge/groove array planforms direct gas flow trappedinside the grooves toward the downstream combustion flow F direction todiscourage gas flow leakage directly from the pressure side of theturbine airfoil toward the suction side of the airfoil in the localizedblade leakage direction L. The forward zone is generally defined betweenthe leading edge and the mid-chord of the blade 92 airfoil at a cutoffpoint where a line parallel to the turbine 80 axis is roughly in tangentto the pressure side surface of the airfoil. From a more gross summaryperspective, the axial length of the forward zone A can also be definedgenerally as roughly one-third to one-half of the total axial length ofthe airfoil. The remainder of the array pattern comprises the aft zoneB. More than two axially oriented planform arrays can be constructed inaccordance with embodiments of the invention. For example forward,middle and aft ridge/groove array planforms can be constructed on theabradable component surface.

The embodiments shown in FIGS. 12-19, 21, 22, 34-35, 37 and 57 havehockey stick-like planform patterns. The forward upstream zone A groovesand ridges are aligned generally parallel (+/−10%) to the combustion gasaxial flow direction F within the turbine 80 (see FIG. 1). The aftdownstream zone B grooves and ridges are angularly oriented opposite theblade rotational direction R. The range of angles is approximately 30%to 120% of the associated turbine blade 92 camber or trailing edgeangle. For design convenience the downstream angle selection can beselected to match any of the turbine blade high or low pressure averaged(linear average line) side wall surface or camber angle (see, e.g.,angle α_(B2) of FIG. 14 on the high pressure side, commencing at thezone B starting surface and ending at the blade trailing edge), thetrailing edge angle (see, e.g., angle α_(B1) of FIG. 15); the anglematching connection between the leading and trailing edges (see, e.g.,angle α_(B1) of FIG. 14); or any angle between such blade geometryestablished angles, such as α_(B3). Hockey stick-like ridge and groovearray planform patterns are as relatively easy to form on an abradablesurface as purely horizontal or diagonal know planform array patterns,but in fluid flow simulations the hockey stick-like patterns have lessblade tip leakage than either of those known unidirectional planformpatterns. The hockey stick-like patterns are formed by knowncutting/abrading or additive layer building methods that have beenpreviously used to form known abradable component ridge and groovepatterns.

In FIG. 12, the abradable component 160 has forward ridges/ridge tips162A/164A and grooves 168A that are oriented at angle α_(A) within +/−10degrees relative to the axial turbine axial flow direction F. The aftridges/ridge tips 162B/164B and grooves 168B are oriented at an angleα_(B) that is approximately the turbine blade 92 trailing edge angle. Asshown schematically in FIG. 12, the forward ridges 162A block theforward zone A blade leakage direction and the rear ridges 162B blockthe aft zone B blade leakage L. Horizontal spacer ridges 169 areperiodically oriented axially across the entire blade 92 footprint andabout the circumference of the abradable component surface 167, in orderto block and disrupt blade tip leakage L, but unlike known design flat,continuous surface abradable surfaces reduce potential surface area thatmay cause blade tip contact and wear.

The abradable component 170 embodiment of FIG. 13 is similar to that ofFIG. 12, with the forward portion ridges 172A/174A and grooves 178Aoriented generally parallel to the turbine combustion gas flow directionF while the rear ridges 172B/174B and grooves 178B are oriented at angleα_(B) that is approximately equal to that formed between the pressureside of the turbine blade 92 starting at zone B to the blade trailingedge. As with the embodiment of FIG. 12, the horizontal spacer ridges179 are periodically oriented axially across the entire blade 92footprint and about the circumference of the abradable component surface167, in order to block and disrupt blade tip leakage L.

The abradable component 180 embodiment of FIG. 14 is similar to that ofFIGS. 12 and 13, with the forward portion ridges 182A/184A and grooves188A oriented generally parallel to the turbine combustion gas flowdirection F while the rear ridges 182B/184B and grooves 188B areselectively oriented at any of angles α_(B1) to α_(B3). Angle α_(B1) isthe angle formed between the leading and trailing edges of blade 92. Asin FIG. 13, angle α_(B2) is approximately parallel to the portion of theturbine blade 92 high pressure side wall that is in opposed relationshipwith the aft zone B. As shown in FIG. 14 the rear ridges 182B/184B andgrooves 188B are actually oriented at angle α_(B3), which is an anglethat is roughly 50% of angle α_(B2). As with the embodiment of FIG. 12,the horizontal spacer ridges 189 are periodically oriented axiallyacross the entire blade 92 footprint and about the circumference of theabradable component surface 187, in order to block and disrupt blade tipleakage L.

In the abradable component 190 embodiment of FIG. 15 the forward ridges192A/194A and grooves 198A and angle α_(A) are similar to those of FIG.14, but the aft ridges 192B/194B and grooves 198B have narrower spacingand widths than FIG. 14. The alternative angle α_(B1) of the aft ridges192B/194B and grooves 198B shown in FIG. 15 matches the trailing edgeangle of the turbine blade 92, as does the angle α_(B) in FIG. 12. Theactual angle α_(B2) is approximately parallel to the portion of theturbine blade 92 high pressure side wall that is in opposed relationshipwith the aft zone B, as in FIG. 13. The alternative angle α_(B3) and thehorizontal spacer ridges 199 match those of FIG. 14, though other arraysof angles or spacer ridges can be utilized.

Alternative spacer ridge patterns are shown in FIGS. 16 and 17. In theembodiment of FIG. 16 the abradable component 200 incorporates an arrayof full-length spacer ridges 209 that span the full axial footprint ofthe turbine blade 92 and additional forward spacer ridges 209A that areinserted between the full-length ridges. The additional forward spacerridges 209A provide for additional blockage or blade tip leakage in theblade 92 portion that is proximal the leading edge. In the embodiment ofFIG. 17 the abradable component 210 has a pattern of full-length spacerridges 219 and also circumferentially staggered arrays of forward spacerridges 219A and aft spacer ridges 219B. The circumferentially staggeredridges 219A/B provide for periodic blocking or disruption of blade tipleakage as the blade 92 sweeps the abradable component 210 surface,without the potential for continuous contact throughout the sweep thatmight cause premature blade tip wear.

While arrays of horizontal spacer ridges have been previously discussed,other embodiments of the invention include vertical spacer ridges. Moreparticularly the abradable component 220 embodiment of FIGS. 18 and 19incorporate forward ridges 222A between which are groove 228A. Thosegrooves are interrupted by staggered forward vertical ridges 223A thatinterconnect with the forward ridges 222A. The vertical As is shown inFIG. 18 the staggered forward vertical ridges 223A form a series ofdiagonal arrays sloping downwardly from left to right. A full-lengthvertical spacer ridge 229 is oriented in a transitional zone T betweenthe forward zone A and the aft zone B. The aft ridges 222B and grooves228B are angularly oriented, completing the hockey stick-like planformarray with the forward ridges 222A and grooves 228A. Staggered rearvertical ridges 223B are arrayed similarly to the forward verticalridges 223A. The vertical ridges 223A/B and 229 disrupt generally axialairflow leakage across the abradable component 220 grooves from theforward to aft portions that otherwise occur with uninterruptedfull-length groove embodiments of FIGS. 12-17, but at the potentialdisadvantage of increased blade tip wear at each potential rubbingcontact point with one of the vertical ridges. Staggered vertical ridges223A/B as a compromise periodically disrupt axial airflow through thegrooves 228A/B without introducing a potential 360 degree rubbingsurface for turbine blade tips. Potential 360 degree rubbing surfacecontact for the continuous vertical ridge 229 can be reduced byshortening that ridge vertical height relative to the ridges 222A/B or223 A/B, but still providing some axial flow disruptive capability inthe transition zone T between the forward grooves 228A and the reargrooves 228B.

FIG. 20 shows a simulated fluid flow comparison between a hockeystick-like ridge/groove pattern array planform with continuous grooves(solid line) and split grooves disrupted by staggered vertical ridges(dotted line). The total blade tip leakage mass flux (area below therespective lines) is lower for the split groove array pattern than forthe continuous groove array pattern.

Staggered ridges that disrupt airflow in grooves do not have to bealigned vertically in the direction of blade rotation R. As shown inFIG. 21 the abradable component 230 has patterns of respective forwardand aft ridges 232A/B and grooves 238A/B that are interrupted by angledpatterns of ridges 233A/B (α_(A), α_(B)) that connect between successiverows of forward and aft ridges and periodically block downstream flowwithin the grooves 238 A/B. As with the embodiment of FIG. 18, theabradable component 230 has a continuous vertically aligned ridge 239located at the transition between the forward zone A and aft zone B. Theintersecting angled array of the ridges 232A and 233A/B effectivelyblock localized blade tip leakage L from the high pressure side 96 tothe low pressure side 98 along the turbine blade axial length from theleading to trailing edges.

It is noted that the spacer ridge 169, 179, 189, 199, 209, 219, 229,239, etc., embodiments shown in FIGS. 12-19 and 21 may have differentrelative heights in the same abradable component array and may differ inheight from one or more of the other ridge arrays within the component.For example if the spacer ridge height is less than the height of otherridges in the abradable surface it may never contact a blade tip but canstill function to disrupt airflow along the adjoining interruptedgroove.

FIG. 22 is an alternative embodiment of a hockey stick-like planformpattern abradable component 240 that combines the embodiment concepts ofdistinct forward zone A and aft zone B respective ridge 242 A/B andgroove 248A/B patterns which intersect at a transition T without anyvertical ridge to split the zones from each other. Thus the grooves248A/B form a continuous composite groove from the leading or forwardedge of the abradable component 240 to its aft most downstream edge (seeflow direction F arrow) that is covered by the axial sweep of acorresponding turbine blade. The staggered vertical ridges 243A/Binterrupt axial flow through each groove without potential continuousabrasion contact between the abradable surface and a correspondingrotating blade (in the direction of rotation arrow R) at one axiallocation. However the relatively long runs of continuous straight-linegrooves 248A/B, interrupted only periodically by small vertical ridges243 A/B, provide for ease of manufacture by water jet erosion or otherknown manufacturing techniques. The abradable component 240 embodimentoffers a good subjective design compromise among airflow performance,blade tip wear and manufacturing ease/cost.

FIGS. 23-25 show embodiments of abradable component ridge and grooveplanform arrays that comprise zig-zag patterns. The zig-zag patterns areformed by adding one or more layers of material on an abradable surfacesubstrate to form ridges or by forming grooves within the substrate,such as by known laser or water jet cutting methods. In FIG. 23 theabradable component 250 substrate surface 257 has a continuous groove258 formed therein, starting at 258′ and terminating at 258″ defines apattern of alternating finger-like interleaving ridges 252. Other grooveand ridge zig-zag patterns may be formed in an abradable component. Asshown in the embodiment of FIG. 24 the abradable component 260 has acontinuous pattern diagonally oriented groove 268 initiated at 268′ andterminating at 268″ formed in the substrate surface 267, leaving angularoriented ridges 262. In FIG. 25 the abradable component embodiment 270has a vee or hockey stick-like dual zone multi groove pattern formed bya pair of grooves 278A and 278B in the substrate surface 277. Groove 278starts at 278′ and terminates at 278″. In order to complete the vee orhockey stick-like pattern on the entire substrate surface 277 the secondgroove 278A is formed in the bottom left hand portion of the abradablecomponent 270, starting at 278A′ and terminating at 278A″. Respectiveblade tip leakage L flow-directing front and rear ridges, 272A and 272B,are formed in the respective forward and aft zones of the abradablesurface 277, as was done with the abradable embodiments of FIGS. 12-19,21 and 22. The groove 258, 268, 278 or 278A do not have to be formedcontinuously and may include blocking ridges like the ridges 223A/B ofthe embodiment of FIGS. 18 and 19, in order to inhibit gas flow throughthe entire axial length of the grooves.

FIGS. 26-29 show embodiments of abradable component ridge and grooveplanform arrays that comprise nested loop patterns. The nested looppatterns are formed by adding one or more layers of material on anabradable surface substrate to form ridges or by forming grooves withinthe substrate, such as by known laser or water jet cutting methods. Theabradable component 280 embodiment of FIG. 26 has an array of verticallyoriented nested loop patterns 281 that are separated by horizontallyoriented spacer ridges 289. Each loop pattern 281 has nested grooves288A-288E and corresponding complementary ridges comprising centralridge 282A loop ridges 282 B-282E. In FIG. 27 the abradable component280′ includes a pattern of nested loops 281A in forward zone A andnested loops 281B in the aft zone B. The nested loops 281A and 281B areseparated by spacer ridges both horizontally 289 and vertically 289A. Inthe abradable embodiment 280″ of FIG. 28 the horizontal portions of thenested loops 281″ are oriented at an angle α. In the abradableembodiment 280″ of FIG. 29 the nested generally horizontal or axialloops 281A″ and 281B′″ are arrayed at respective angles α_(A) and α_(B)in separate forward zone A and aft zone B arrays. The fore and aftangles and loop dimensions may be varied to minimize blade tip leakagein each of the zones.

FIGS. 30-33 show embodiments of abradable component ridge and grooveplanform arrays that comprise spiral maze patterns, similar to thenested loop patterns. The maze patterns are formed by adding one or morelayers of material on an abradable surface substrate to form ridges.Alternatively, as shown in these related figures, the maze pattern iscreated by forming grooves within the substrate, such as by known laseror water jet cutting methods. The abradable component 290 embodiment ofFIG. 30 has an array of vertically oriented nested maze patterns 291,each initiating at 291A and terminating at 291B, that are separated byhorizontally oriented spacer ridges 299. In FIG. 31 the abradablecomponent 290′includes a pattern of nested mazes 291A in forward zone Aand nested mazes 291B in the aft zone B. The nested mazes 291A and 291Bare separated by spacer ridges both horizontally 299′ and vertically293′. In the abradable embodiment 290″ of FIG. 32 the horizontalportions of the nested mazes 291″ are oriented at an angle α. In theabradable embodiment 290′″ of FIG. 33 the generally horizontal portionsof mazes 291A′″ and 291B′″ are arrayed at respective angles α_(A) andα_(B) in separate forward zone A and aft zone B arrays, while thegenerally vertical portions are aligned with the blade rotational sweep.The fore and aft angles α_(A) and α_(B) and maze dimensions may bevaried to minimize blade tip leakage in each of the zones.

FIGS. 34 and 35 are directed to an abradable component 300 embodimentwith separate and distinct multi-arrayed ridge 302A/302B and groove308A/308B pattern in the respective forward zone A and aft zone B thatare joined by a pattern of corresponding curved ridges 302T and grooves308T in a transition zone T. In this exemplary embodiment pattern thegrooves 308A/B/T are formed as closed loops within the abradablecomponent 300 surface, circumscribing the corresponding ribs 302A/B/T.Inter-rib spacing S_(RA), S_(RB) and S_(RT) and corresponding groovespacing may vary axially and vertically across the component surface inorder to minimize local blade tip leakage. As will be described ingreater detail herein, rib and groove cross sectional profile may beasymmetrical and formed at different angles relative to the abradablecomponent 300 surface in order to reduce localized blade tip leakage.FIG. 36 shows comparative fluid dynamics simulations of comparable depthridge and groove profiles in abradable components. The solid linerepresents blade tip leakage in an abradable component of the type ofFIGS. 34 and 35. The dashed line represents a prior art type abradablecomponent surface having only axial or horizontally oriented ribs andgrooves. The dotted line represents a prior art abradable componentsimilar to that of FIG. 7 with only diagonally oriented ribs and groovesaligned with the trailing edge angle of the corresponding turbine blade92. The abradable component 300 had less blade tip leakage than theleakage of either of the known prior art type unidirectional abradablesurface ridge and groove patterns.

Abradable Surface Ridge and Groove Cross Sectional Profiles

Exemplary invention embodiment abradable surface ridge and groove crosssectional profiles are shown in FIGS. 37 41 and 43 63. Unlike knownabradable cross sectional profile patterns that have uniform heightacross an entire abradable surface, many of the present invention crosssectional profiles formed in the thermally sprayed abradable layercomprise composite multi height/depth ridge and groove patterns thathave distinct upper (zone I) and lower (zone II) wear zones. The lowerzone II optimizes engine airflow and structural characteristics whilethe upper zone I minimizes blade tip gap and wear by being more easilyabradable than the lower zone. Various embodiments of the abradablecomponent afford easier abradability of the upper zone with upper subridges or nibs having smaller cross sectional area than the lower zonerib structure. In some embodiments the upper sub ridges or nibs areformed to bend or otherwise flex in the event of minor blade tip contactand wear down and/or shear off in the event of greater blade tipcontact. In other embodiments the upper zone sub ridges or nibs arepixelated into arrays of upper wear zones so that only those nibs inlocalized contact with one or more blade tips are worn while othersoutside the localized wear zone remain intact. While upper zone portionsof the ridges are worn away they cause less blade tip wear than priorknown monolithic ridges and afford greater profile forming flexibilitythan CMC/FGI abradable component constructions that require profilingaround the physical constraints of the composite hollow ceramic spherematrix orientations and diameters. In embodiments of the invention asthe upper zone ridge portion is worn away the remaining lower ridgeportion preserves engine efficiency by controlling blade tip leakage. Inthe event that the localized blade tip gap is further reduced, the bladetips wear away the lower ridge portion at that location. However therelatively higher ridges outside that lower ridge portion localized weararea maintain smaller blade tip gaps to preserve engine performanceefficiency.

With the progressive wear zones construction of some embodiments of theinvention blade tip gap G can be reduced from previously acceptableknown dimensions. For example, if a known acceptable blade gap G designspecification is 1 mm the higher ridges in wear zone I can be increasedin height so that the blade tip gap is reduced to 0.5 mm. The lowerridges that establish the boundary for wear zone II are set at a heightso that their distal tip portions are spaced 1 mm from the blade tip. Inthis manner a 50% tighter blade tip gap G is established for routineturbine operation, with acceptance of some potential wear caused byblade contact with the upper ridges in zone I. Continued localizedprogressive blade wearing in zone II will only be initiated if the bladetip encroaches into the lower zone, but in any event the blade tip gap Gof 1 mm is no worse than known blade tip gap specifications. In someexemplary embodiments the upper zone I height is approximately ⅓ to ⅔ ofthe lower zone II height.

The abradable component 310 of FIGS. 37-41 has alternating height curvedridges 312A and 312B that project up from the abradable surface 317 andstructurally supported by the support surface 311. Grooves 318 separatethe alternating height ridges 312A/B and are defined by the ridge sidewalls 315A/B and 316A/B. Wear zone I is established from the respectivetips 314A of taller ridges 312A down to the respective tips 314B of thelower ridges 312B. Wear zone II is established from the tips 314B downto the substrate surface 317. Under turbine operating conditions (FIGS.39 and 40) the blade gap G is maintained between the higher ridge tips312A and the blade tip 94. While the blade gap G is maintained bladeleakage L travels in the blade 92 rotational direction (arrow R) fromthe higher pressurized side of the blade 96 (at pressure P_(P)) to thelow or suction pressurized side of the blade 98 (at pressure P_(S)).Blade leakage L under the blade tip 94 is partially trapped between anopposed pair of higher ridges 312A and the intermediate lower ridge312B, forming a blocking swirling pattern that further resists the bladeleakage. If the blade tip gap G becomes reduced for any one or moreblades due to turbine casing 100 distortion, fast engine startup mode orother reason initial contact between the blade tip 94 and the abradablecomponent 310 will occur at the higher ridge tips 314A. While still inzone I the blade tips 94 only rub the alternate staggered higher ridges312A. If the blade gap G progressively becomes smaller, the higherridges 312A will be abraded until they are worn all the way through zoneI and start to contact the lower ridge tips 314B in zone II. Once inZone II the turbine blade tip 94 rubs all of the remaining ridges 314A/Bat the localized wear zone, but in other localized portions of theturbine casing there may be no reduction in the blade tip gap G and theupper ridges 312 A may be intact at their full height. Thus thealternating height rib construction of the abradable component 310accommodates localized wear within zones I and II, but preserves theblade tip gap G and the aerodynamic control of blade tip leakage L inthose localized areas where there is no turbine casing 100 or blade 92distortion. When either standard or fast start or both engine operationmodes are desired the taller ridges 312A form the primary layer ofclearance, with the smallest blade tip gap G, providing the best energyefficiency clearance for machines that typically utilize lower ramprates or that do not perform warm starts. Generally the ridge heightH_(RB) for the lower ridge tips 314B is between 25%-75% of the higherridge tip 314A height, H_(RA). In the embodiment shown in FIG. 41 thecenterline spacing S_(RA) between successive higher ridges 312A equalsthe centerline spacing S_(RB) between successive lower ridges 312B.Other centerline spacing and patterns of multi height ridges, includingmore than two ridge heights, can be employed.

Other embodiments of ridge and groove profiles with upper and lower wearzones include the stepped ridge profiles of FIGS. 43 and 44, which arecompared to the known single height ridge structure of the prior artabradable 150 in FIG. 42. Known single height ridge abradables 150include a base support 151 that is coupled to a turbine casing 100, asubstrate surface 157 and symmetrical ridges 152 having inwardly slopingside walls 155, 156 that terminate in a flat ridge tip 154. The ridgetips 154 have a common height and establish the blade tip gap G with theopposed, spaced blade tip 94. Grooves 158 are established between ridges152. Ridge spacing S_(R), groove width W_(G) and ridge width W_(R) areselected for a specific application. In comparison, the stepped ridgeprofiles of FIGS. 43 and 44 employ two distinct upper and lower wearzones on a ridge structure.

The abradable component 320 of FIG. 43 has a support surface 321 and anabradable surface 327 upon which are arrayed distinct two-tier ridges:lower ridge 322B and upper ridge 322A. The lower ridge 322B has a pairof sidewalls 325B and 326B that terminate in plateau 324B of heightH_(RB). The upper ridge 322A is formed on and projects from the plateau324B, having side walls 325A and 326A terminating in a distal ridge tip324A of height H_(RA) and width W_(R). The ridge tip 324A establishesthe blade tip gap G with an opposed, spaced blade tip 94. Wear zone IIextends vertically from the abradable surface 327 to the plateau 324Band wear zone I extends vertically from the plateau 324B to the ridgetip 324A. The two rightmost ridges 322A/B in FIG. 43 have asymmetricalprofiles with merged common side walls 326A/B, while the oppositesidewalls 325A and 325B are laterally offset from each other andseparated by the plateau 324B of width W. Grooves 328 are definedbetween the ridges 322A/B. The leftmost ridge 322A′/B′ has a symmetricalprofile. The lower ridge 322B′ has a pair of converging sidewalls 325B′and 326B′, terminating in plateau 324B′. The upper ridge 322A′ iscentered on the plateau 324B′, leaving an equal width offset W_(P′) withrespect to the upper ridge sidewalls 325A′ and 326A′. The upper ridgetip 324A′ has width W_(R′). Ridge spacing S_(R) and groove width W_(G)are selected to provide desired blade tip leakage airflow control. Insome exemplary embodiments of abradable component ridge and grooveprofiles described herein the groove widths W_(G) are approximately ⅓-⅔of lower ridge width. While the ridges and grooves shown in FIG. 43 aresymmetrically spaced, other spacing profiles may be chosen, includingdifferent ridge cross sectional profiles that create the stepped wearzones I and II.

FIG. 44 shows another stepped profile abradable component 330 with theridges 332A/B having vertically oriented parallel side walls 335A/B and336A/B. The lower ridge terminates in ridge plateau 334B, upon which theupper ridge 332A is oriented and terminates in ridge tip 334A. In someapplications it may be desirable to employ the vertically orientedsidewalls and flat tips/plateaus that define sharp-cornered profiles,for airflow control in the blade tip gap. The upper wear zone I isbetween the ridge tip 334A and the ridge plateau 334B and the lower wearzone is between the plateau and the abradable surface 337. As with theabradable embodiment 320 of FIG. 43, while the ridges and grooves shownin FIG. 44 are symmetrically spaced, other spacing profiles may bechosen, including different ridge cross sectional profiles that createthe stepped wear zones I and II.

In another permutation or species of stepped ridge constructionabradable components, separate upper and lower wear zones I and II alsomay be created by employing multiple groove depths, groove widths andridge widths, as employed in the abradable 340 profile shown in FIG. 45.The lower rib 342B has rib plateau 344B that defines wear zone II inconjunction with the abradable surface 347. The rib plateau 344Bsupports a pair of opposed, laterally flanking upper ribs 342A, whichterminate in common height rib tips 344A. The wear zone I is definedbetween the rib tips 344A and the plateau 344B. A convenient way to formthe abradable component 340 profiles is to cut dual depth grooves 348Aand 348B into a flat surfaced abradable substrate at respective depthsD_(GA) and D_(GB). Ridge spacing S_(R), groove width W_(GA/B) and ridgetip 344A width W_(R) are selected to provide desired blade tip leakageairflow control. While the ridges and grooves shown in FIG. 45 aresymmetrically spaced, other spacing profiles may be chosen, includingdifferent ridge cross sectional profiles that create the stepped wearzones I and II.

As shown in FIG. 46, in certain turbine applications it may be desirableto control blade tip leakage by employing an abradable component 350embodiment having asymmetric profile abradable ridges 352 withvertically oriented, sharp-edged upstream sidewalls 356 and slopingopposite downstream sidewalls 355 extending from the substrate surface357 and terminating in ridge tips 354. Blade leakage L is initiallyopposed by the vertical sidewall 356. Some leakage airflow L nonethelessis compressed between the ridge tip 354 and the opposing blade tip 94while flowing from the high pressure blade side 96 to the lower pressuresuction blade side 98 of the blade. That leakage flow follows thedownward sloping ridge wall 355, where it is redirected opposite bladerotation direction R by the vertical sidewall 356 of the next downstreamridge. The now counter flowing leakage air L opposes further incomingleakage airflow L in the direction of blade rotation R. Dimensionalreferences shown in FIG. 46 are consistent with the referencedescriptions of previously described figures. While the abradablecomponent embodiment 350 of FIG. 46 does not employ the progressive wearzones I and II of other previously described abradable componentprofiles, such zones may be incorporated in other below-describedasymmetric profile rib embodiments.

Progressive wear zones can be incorporated in asymmetric ribs or anyother rib profile by cutting grooves into the ribs, so that remainingupstanding rib material flanking the groove cut has a smaller horizontalcross sectional area than the remaining underlying rib. Grooveorientation and profile may also be tailored to enhance airflowcharacteristics of the turbine engine by reducing undesirable blade tipleakage, is shown in the embodiment of FIG. 47 to be describedsubsequently herein. In this manner, the thermally sprayed abradablecomponent surface is constructed with both enhanced airflowcharacteristics and reduced potential blade tip wear, as the blade tiponly contacts portions of the easier to abrade upper wear zone I. Thelower wear zone II remains in the lower rib structure below the groovedepth. Other exemplary embodiments of abradable component ridge andgroove profiles used to form progressive wear zones are now described.Structural features and component dimensional references in theseadditional embodiments that are common to previously describedembodiments are identified with similar series of reference numbers andsymbols without further detailed description.

FIG. 47 shows an abradable component 360 having the rib cross sectionalprofile of the FIG. 46 abradable component 350, but with inclusion ofdual level grooves 368A formed in the ridge tips 364 and 368B formedbetween the ridges 362 to the substrate surface 367. The upper grooves368A form shallower depth D_(G) lateral ridges that comprise the wearzone I while the remainder of the ridge 362 below the groove depthcomprises the lower wear zone II. In this abradable component embodiment360 the upper grooves 368A are oriented parallel to the ridge 362longitudinal axis and are normal to the ridge tip 364 surface, but othergroove orientations, profiles and depths may be employed to optimizeairflow control and/or minimize blade tip wear.

In the abradable component 370 embodiment of FIG. 48 a plurality ofupper grooves 378A are tilted fore-aft relative to the ridge tip 374 atangle γ, depth D_(GA) and have parallel groove side walls. Upper wearzone I is established between the bottom of the groove 378A and theridge tip 374 and lower wear zone II is below the upper wear zone downto the substrate surface 377. In the alternative embodiment of FIG. 49the abradable component 380 has upper grooves 388A with rectangularprofiles that are skewed at angle A relative to the ridge 382longitudinal axis and its sidewalls 385/386. The upper groove 388A asshown is also normal to the ridge tip 384 surface. The upper wear zone Iis above the groove depth D_(GA) and wear zone II is below that groovedepth down to the substrate surface 387. For brevity the remainder ofthe structural features and dimensions are labelled in FIGS. 48 and 49with the same conventions as the previously described abradable surfaceprofile embodiments and has the same previously described functions,purposes and relationships.

As shown in FIGS. 50-52, upper grooves do not have to have parallelsidewalls and may be oriented at different angles relative to the ridgetip surface. Also upper grooves may be utilized in ridges having variedcross sectional profiles. The ridges of the abradable componentembodiments 390, 400 and 410 have symmetrical sidewalls that converge ina ridge tip. As in previously described embodiments having dual heightgrooves, the respective upper wear zones I are from the ridge tip to thebottom of the groove depth D_(G) and the lower wears zones II are fromthe groove bottom to the substrate surface. In FIG. 50 the upper groove398A is normal to the substrate surface (ε=90°) and the groove sidewallsdiverge at angle Φ. In FIG. 51 the groove 408A is tilted at angle +εrelative to the substrate surface and the groove 418A in FIG. 52 istilted at −ε relative to the substrate surface. In both of the abradablecomponent embodiments 400 and 410 the upper groove sidewalls diverge atangle Φ. For brevity the remainder of the structural features anddimensions are labelled in FIGS. 50-52 with the same conventions as thepreviously described abradable surface profile embodiments and has thesame previously described functions, purposes and relationships.

In FIGS. 53-56 the abradable ridge embodiments shown have trapezoidalcross sectional profiles and ridge tips with upper grooves in variousorientations, for selective airflow control, while also having selectiveupper and lower wear zones. In FIG. 53 the abradable component 430embodiment has an array of ridges 432 with asymmetric cross sectionalprofiles, separated by lower grooves 438B. Each ridge 432 has a firstside wall 435 sloping at angle β₁ and a second side wall 436 sloping atangle β₂. Each ridge 432 has an upper groove 438A that is parallel tothe ridge longitudinal axis and normal to the ridge tip 434. The depthof upper groove 438A defines the lower limit of the upper wear zone Iand the remaining height of the ridge 432 defines the lower wear zoneII.

In FIGS. 54-56 the respective ridge 422, 442 and 452 cross sections aretrapezoidal with parallel side walls 425/445/455 and 426/446/456 thatare oriented at angle β. The right side walls 426/446/456 are orientedto lean opposite the blade rotation direction, so that air trappedwithin an intermediate lower groove 428B/448B/458B between two adjacentridges is also redirected opposite the blade rotation direction,opposing the blade tip leakage direction from the upstream high pressureside 96 of the turbine blade to the low pressure suction side 98 of theturbine blade, as was shown and described in the asymmetric abradableprofile 350 of FIG. 46. Respective upper groove 428A/448A/458Aorientation and profile are also altered to direct airflow leakage andto form the upper wear zone I. Groove profiles are selectively alteredin a range from parallel sidewalls with no divergence to negative orpositive divergence of angle Φ, of varying depths D_(G) and at varyingangular orientations ε with respect to the ridge tip surface. In FIG. 54the upper groove 428A is oriented normal to the ridge tip 424 surface(ε=90°). In FIGS. 55 and 56 the respective upper grooves 448A and 458Aare oriented at angles +/−ε with respect its corresponding ridge tipsurface.

FIG. 57 shows an abradable component 460 planform incorporatingmulti-level grooves and upper/lower wear zones, with forward A and aft Bridges 462A/462B separated by lower grooves 468A/B that are oriented atrespective angles α_(A/B). Arrays of fore and aft upper partial depthgrooves 463A/B of the type shown in the embodiment of FIG. 49 are formedin the respective arrays of ridges 462A/B and are oriented transversethe ridges and the full depth grooves 468A/B at respective anglesβ_(A/B). The upper partial depth grooves 463A/B define the verticalboundaries of the abradable component 460 upper wear zones I, with theremaining portions of the ridges below those partial depth upper groovesdefining the vertical boundaries of the lower wear zones II.

With thermally sprayed abradable component construction, the crosssections and heights of upper wear zone I thermally sprayed abradablematerial can be configured to conform to different degrees of blade tipintrusion by defining arrays of micro ribs or nibs, as shown in FIG. 58,on top of ridges, without the aforementioned geometric limitations offorming grooves around hollow ceramic spheres in CMC/FGI abradablecomponent constructions, and the design benefits of using a metallicabradable component support structure. The abradable component 470includes a previously described metallic support surface 471, witharrays of lower grooves and ridges forming a lower wear zone II.Specifically the lower ridge 472B has side walls 475B and 476B thatterminate in a ridge plateau 474B. Lower grooves 478B are defined by theridge side walls 475B and 476B and the substrate surface 477. Micro ribsor nibs 472A are formed on the lower ridge plateau 474B by knownadditive processes or by forming an array of intersecting grooves 478Aand 478C within the lower ridge 472B, without any hollow sphereintegrity preservation geometric constraints that would otherwise beimposed in a CMC/FGI abradable component design. In the embodiment ofFIG. 58 the nibs 472A have square or other rectangular cross section,defined by upstanding side walls 475A, 475C, 476A and 476C thatterminate in ridge tips 474A of common height. Other nib 472A crosssectional planform shapes can be utilized, including by way of exampletrapezoidal or hexagonal cross sections. Nib arrays including differentlocalized cross sections and heights can also be utilized.

In the alternative embodiment of FIG. 60, distal rib tips 474A′ of theupstanding pixelated nib 472A′ are constructed of thermally sprayedmaterial 480 having different physical properties and/or compositionsthan the lower thermally sprayed material 482. For example, the upperdistal material 480 can be constructed with easier or less abrasiveabrasion properties (e.g., softer or more porous or both) than the lowermaterial 482. In this manner the blade tip gap G can be designed to beless than used in previously known abradable components to reduce bladetip leakage, so that any localized blade intrusion into the material 480is less likely to wear the blade tips, even though such contact becomesmore likely. In this manner the turbine engine can be designed withsmaller blade tip gap, increasing its operational efficiency, as well asits ability to be operated in standard or fast start startup mode, whilenot significantly impacting blade wear.

Nib 472A and groove 478A/C dimensional boundaries are identified inFIGS. 58 and 59, consistent with those described in the priorembodiments. Generally nib 472A height H ranges from approximately20%400% of the blade tip gap G or from approximately ⅓-⅔ the total ridgeheight of the lower ridge 472B and the nibs 472A. Nib 472A cross sectionranges from approximately 20% to 50% of the nib height H_(BA). Nibmaterial construction and surface density (quantified by centerlinespacing S_(RA/B) and groove width W_(GA)) are chosen to balanceabradable component 470 wear resistance, thermal resistance, andstructural stability and airflow characteristics. For example, aplurality of small width nibs 472A produced in a controlled densitythermally sprayed ceramic abradable offers high leakage protection tohot gas. These can be at high incursion prone areas only or the fullengine set. It is suggested that were additional sealing is needed thisis done via the increase of plurality of the ridges maintaining theirlow strength and not by increasing the width of the ridges. Typical nibcenterline spacing S_(RA/B) or nib 472A structure and array patterndensity selection enables the pixelated nibs to respond in differentmodes to varying depths of blade tip 94 incursions, as shown in FIGS.61-63.

In FIG. 61 there is no or actually negative blade tip gap G, as theturbine blade tip 94 is contacting the ridge tips 474A of the pixelatednibs 472A. The blade tip 94 contact intrusion flexes the pixelated nibs472A. In FIG. 62 there is deeper blade tip intrusion into the abradablecomponent 470, causing the nibs 472A to wear, fracture or shear off thelower rib plateau 474B, leaving a residual blade tip gap there between.In this manner there is minimal blade tip contact with the residualbroken nib stubs 472A (if any), while the lower ridge 472B in wear zoneII maintains airflow control of blade tip leakage. In FIG. 63 the bladetip 94 has intruded into the lower ridge plateau 474B of the lower rib472B in wear zone II. Returning to the example of engines capable ofstartup in either standard or fast start mode, in an alternativeembodiment the nibs 472A can be arrayed in alternating height H_(RA)patterns: the higher optimized for standard startup and the loweroptimized for fast startup. In fast startup mode the higher of thealternating nibs 472A fracture, leaving the lower of the alternatingnibs for maintenance of blade tip gap G. Exemplary thermally sprayedabradable components having frangible ribs or nibs have height H_(RA) towidth W_(RA) ratio of greater than 1. Typically the width W_(RA)measured at the peak of the ridge or nib would be 0.5-2 mm and itsheight H_(RA) is determined by the engine incursion needs and maintain aheight to width ratio (H_(RA)/W_(RA)) greater than 1. It is suggestedthat where additional sealing is needed, this is done via the increaseof plurality of the ridges or nibs (i.e., a larger distribution density,of narrow width nibs or ridges, maintaining their low strength) and notby increasing their width W_(RA). For zones in the engine that requirethe low speed abradable systems the ratio of ridge or nib widths togroove width (W_(RA)/W_(GA)) is preferably less than 1. For engineabradable component surface zones or areas that are not typically inneed of easy blade tip abradability, the abradable surface crosssectional profile is preferably maximized for aerodynamic sealingcapability (e.g., small blade tip gap G and minimized blade tip leakageby applying the surface planform and cross sectional profile embodimentsof the invention, with the ridge/nib to groove width ratio of greaterthan 1.

Multiple modes of blade depth intrusion into the circumferentialabradable surface may occur in any turbine engine at differentlocations. Therefore, the abradable surface construction at anylocalized circumferential position may be varied selectively tocompensate for likely degrees of blade intrusion. For example, referringback to the typical known circumferential wear zone patterns of gasturbine engines 80 in FIGS. 3-6, the blade tip gap G at the 3:00 and6:00 positions may be smaller than those wear patterns of the 12:00 and9:00 circumferential positions. Anticipating greater wear at the 12:00and 6:00 positions the lower ridge height H_(RB) can be selected toestablish a worst-case minimal blade tip gap G and the pixelated orother upper wear zone I ridge structure height H_(RA), cross sectionalwidth, and nib spacing density can be chosen to establish a small “bestcase” blade tip gap G in other circumferential positions about theturbine casing where there is less or minimal likelihood abradablecomponent and case distortion that might cause the blade tip 94 tointrude into the abradable surface layer. Using the frangible ridges472A of FIG. 62 as an example, during severe engine operating conditions(e.g. when the engine is in fast start startup mode) the blade 94impacts the frangible ridges 472A or 472A′—the ridges fracture under thehigh load increasing clearance at the impact zones only—limiting theblade tip wear at non optimal abradable conditions. Generally, the upperwear zone I ridge height in the abradable component can be chosen sothat the ideal blade tip gap is 0.25 mm. The 3:00 and 9:00 turbinecasing circumferential wear zones (e.g., 124 and 128 of FIG. 6) arelikely to maintain the desired 0.25 mm blade tip gap throughout theengine operational cycles, but there is greater likelihood of turbinecasing/abradable component distortion at other circumferentialpositions. The lower ridge height may be selected to set its ridge tipat an idealized blade tip gap of 1.0 mm so that in the higher wear zonesthe blade tip only wears deeper into the wear zone I and never contactsthe lower ridge tip that sets the boundary for the lower wear zone II.If despite best calculations the blade tip continues to wear into thewear zone II, the resultant blade tip wear operational conditions are noworse than in previously known abradable layer constructions. However inthe remainder of the localized circumferential positions about theabradable layer the turbine is successfully operating with a lower bladetip gap G and thus at higher operational efficiency, with little or noadverse increased wear on the blade tips.

Embodiments Including Pixelated Major Planform Patterns (PMPP) ofDiscontinuous Micro Surface Features (MSF)

Embodiments described herein can be readily utilized in abradablecomponents for turbine engines, including gas turbine engines. Invarious embodiments, the abradable component includes a support surfacefor coupling to a turbine casing and a thermally sprayedceramic/metallic abradable substrate coupled to the support surface fororientation proximal a rotating turbine blade tip circumferential sweptpath. An elongated pixelated major planform pattern (PMPP) comprising aplurality of discontinuous micro surface features (MSF) project from thesubstrate surface across a majority of the circumferential swept pathfrom a tip to a tail of the turbine blade. In some exemplary embodimentsthe PMPP aggregate planform mimics the general planform of solidprotruding rib abradable components, such as curved or diagonal knowndesigns. In other exemplary embodiments the PMPP aggregate planformmimics the inventive rib and groove planform, hockey stick-like,zig-zag, nested loop, maze and varying curve embodiments shown anddescribed herein. The PMPP repeats radially along the swept path in theblade tip rotational direction, for selectively directing airflowbetween the blade tip and the substrate surface. Each MSF is defined bya pair of first opposed lateral walls defining a width, length andheight that occupy a volume envelope of 1-12 cubic millimeters. In someembodiments the ratio of MSF length and gap defined between each MSF isin the range of approximately 1:1 to 1:3. In other embodiments theration of MSF width and gap is in the range of approximately 1:3 to 1:5.In some embodiment the ratio of MSF height to width is approximately 0.5to 1.0. Feature dimensions can be (but not limited to) between 1 mm and3 mm, with a wall height of between 0.1 mm to 2 mm and a wall thicknessof between 0.2 mm and 1 mm. In some embodiments the PMPP has firstheight and higher second height MSFs.

Either the MSFs in the PMPPs of some embodiments are generated from acast in or an engineered surface feature formed directly in thesubstrate material. In other embodiments the MSFs in the PMPPs aregenerated in the substrate or in an overlying bond coat (BC) layer by anablative or additive surface modification technique such as water jet orelectron beam or laser cutting or by laser sintering methods. Theengineered surface feature will then be coated with high temperatureabradable thermal barrier coating (TBC), with or without an intermediatebond coat layer applied on the engineered MSF features in the PMPP, toproduce a discontinuous surface that will abrade more efficiently than acurrent state of the art coating. Once contacted (by a passing bladetip), released (abraded) particles are removed via a tortuous,convoluted (above or subsurface) path in gaps between the MSFs oradditional slots formed within the abradable surface between the MSFs.Optional continuous slots and/or gaps are oriented so as to provide atortuous path for hot gas ejection, thereby maintaining the sealingefficiency of the primary (contact) surface. The surface configuration,which reduces potential rubbing contact surface area between the bladetips and the discontinuous MSFs, reduces frictional heat generated inthe blade tip. Reduced frictional heat in the blade tip potentiallyreduces worn blade tip material loss attributable to tip over heatingand metal smear/transfer onto the surface of the abradable. Furtherbenefits include the ability to deposit thicker, more robust thermalbarrier coatings over the MSFs than normally possible with knowncontinuous abradable rib designs, thereby imparting potentially extendeddesign life for ring segments.

The abradable embodiments of the invention, which comprise PMPPengineered features with discontinuous MSFs, facilitate optimization ofpotential blade rubbing surface area, optimized angle and planform ofthe PMPPs for guiding airflow in the abradable surface/blade tip gap andoptimized underlying flow/ejection path for abraded particles generatedduring abradable/blade tip rubbing. The micro surface feature (MSF) inits simplest form can be basic shape geometry, repeated in unit cellsacross the surface of the ring segment with gaps between respectivecells. The unit cell MSFs are analogous to pixels that in aggregateforms the PMPP's larger pattern. In more optimized forms the MSF can bemodified according to the requirement of the blade tip relationship ofthe thermal behavior of the component during operation. In suchcircumstances, feature depth, orientation, angle and aspect ratio may bemodified within the surface to produce optimized abradable performancefrom beginning to end of blade sweep. Other optimization parametersinclude ability of thermal spray equipment that forms the TBC topenetrate fully captive areas within the surface and allow for aneffective continuous TBC coating across the entire surface.

As previously noted, the abradable component with the PMPPs comprisingarrays of MSFs is formed by casting the MSFs directly into the abradablesubstrate during its manufacture or by additive manufacturingtechniques, such as electron beam or laser beam deposition, or byablation of substrate material. In the first-noted formation process, asurface feature can be formed in a wax pattern, which is then shelledand cast per standardized investment casting procedures. Alternatively,a ceramic shell insert can be used on the outside of the wax pattern toform part of the shell structure. When utilizing a ceramic shell insertthe MSFs can be more effectively protected during the abradablecomponent manufacture handing and also can more exotic in feature shapeand geometry (i.e., can contain undercuts or fragile protruding featuresthat would not survive a normal shelling operation.

MSFs can be staggered (stepped) to accept and specifically deflectplasma splats for optimum TBC penetration. Surface features cast-in anddeposited onto the substrate may not necessarily fully translate in formto a fully TBC coated surface. During coating, ceramic deposition willbuild upon the substrate in a generally transformative nature but willnot directly duplicate the original engineered surface feature. Thethermal spray thickness can also be a factor in determining finalsurface form. Generally, the thicker the thermal spray coating, the moredissipated the final surface geometry. This is not necessarilyproblematical but needs to be taking into consideration when designingthe engineered surface feature (both initial size and aspect ratio. Forexample, a chevron-shaped MSF formed in the substrate, when subsequentlycoated by an intermediate bond coat layer and a TBC top layer maydissipate as a crescent- or mount-shaped protrusion in the finishedabradable surface projecting profile.

Where exemplary MSF unit cells are shown in FIGS. 64-83, these areprovided for dimensional considerations. For effective dimensionalguidance, the unit cell size can be considered a cube ranging from 1 mmto 12 mm in size. Variations on the cube dimensions can also be appliedto cell height. This can be either smaller or larger than the cube sizedepending upon the geometry of the feature and the thickness of coatingto be applied. Typically the size range of this dimension can be between1 mm and 10 mm.

Various exemplary embodiments described herein, which incorporatepixelated major planform patterns (PMPP) of discontinuous micro surfacefeatures (MSF) jointly or severally in different combinations have atleast some of the following features:

The PMPPs comprising MSF engineered surface features create anunderlying surface with a raised, discontinuous coated structure thatresults in a reduced surface area that is abraded by a passing bladetip.

The MSF engineered surface features improve the adhesion and mechanicalinterlocking properties of the plasma sprayed the abradable coating, dueto increased bonding surface area and the uniqueness of the surfacefeatures to interlock the coating normal to the surface via variousinterlocking geometries that have been described herein.

The engineered micro surface feature (MSF), by virtue of its underlyingaverage surface depth, results in an aggregately thicker coating thatimproves thermal protection for the underlying substrate, leading topotentially cooler substrate temperature.

Due to reduced abradable surface contact area with turbine blade tips,relatively more expensive coatings that are more abradable than standardcost 8YSZ thermal barrier coating material, such as 33YBZO (33%Yb₂O₃—Zirconia) or Talon-type YSZ (high porosity YSZ co-sprayed withpolymer) are not needed. The less abradable (i.e., harder) YSZ wearingof blade tips is negated by the smaller surface area potential rubbingcontact with the rotating blade tips.

The micro surface features (MSF)—some as small as 100 μm inheight—reduce potential thermal barrier coating spallation, due to theincreased adhesion surface contact area with the overlying thermalbarrier coating.

Exemplary embodiments of turbine abradable components includingpixelated major planform patterns (PMPP) of discontinuous micro surfacefeatures (MSF) are shown in FIGS. 64-83. For drawing simplicity theFIGS. 64-66 show schematically PMPPs comprising two rows of MSFs.However, one or more of the PMPPs in any abradable component cancomprise a single row or more than two rows of MSFs. For example, FIG.64 is a planform schematic view of an abradable component 500 split intoupper and lower portions, having a metallic substrate 501. On the upperportion above the split the substrate 501 has a curved overall profilepixelated major planform pattern (PMPP) 502 comprising an array ofchevron-shaped micro surface features (MSF) 503 formed directly on thesubstrate. As previously described the MSFs 503 are formed by any one ormore of a casting process that directly creates them during thesubstrate initial formation; an additive process, building MSFs on thepreviously formed substrate 501 surface; or by an ablative process thatcuts or removes metal from the substrate, leaving the formed MSFs in theremaining material.

On the uppermost portion of the abradable component 500 a thermalbarrier coating (TBC) 506 has been applied directly over the MSFs 503,leaving mound or crescent-shaped profile projections on the abradablecomponent in a PMPP 502 that are arrayed for directing hot gas flowbetween the abradable component and a rotating turbine blade tip. In theevent of contact between the blade tip and the opposing surface of theabradable component 500 the relatively small cross sectional surfacearea MSFs 503 will rub against and be abraded by the blade tip. The MSF503 and turbine blade tip contact is less likely to cause blade tiperosion or abradable 500 surface spallation from the contact compared topreviously known continuous rib or solid surface abradable components,such as those shown in FIGS. 3-11.

On the lowermost portion of the abradable component 500 a metallic bondcoat (BC) 504 is applied to the substrate 501 and the chevron-shapedMSFs 505 are formed in the BC by additive or ablative manufacturingprocesses. The BC 504 and the MSFs 505, arrayed in the PMPP 502, arethen covered with a TBC 506 leaving generally chevron-shaped MSFs 508that project from the substrate 500 surface.

An alternate embodiment abradable component 510 is shown in FIG. 65,wherein the diagonal planform PMPPs 512 are formed in the BC 514 andcomprise arrays of chevron-shaped MSFs 515. The BC 514 and its MSFs 515are then covered with TBC 516 leaving crescent-shaped MSFs 517projecting from the substrate 510 exposed surface. The PMPPs 512 have adiagonal orientation similar to that of the known abradable component130 of FIG. 7.

FIG. 66 is an abradable surface 520 having hockey stick-like PMPP arrayprofiles 522 that are similar to the rib planform patterns of theembodiments of FIGS. 12-22. In the abradable component 520 micro surfacefeatures (MSF) 523 are formed in the substrate surface 521. A bond coat524 is applied on the existing MSFs 522 previously formed in thesubstrate 501 (e.g., by thermal spray coating), leaving more pronouncedand higher MSFs 525. The TBC 526 is applied over the MSFs 522 and the BC524, leaving higher mounded crescent-shaped MSFs 527.

In FIGS. 67 and 68 the abradable component 530 has on its top surface531 discontinuous surface feature PMPPs comprising a seven rowherringbone-like pattern of alternating erect and invertedchevron-shaped MSFs 532, having closed continuous leading edges 533,trailing edges 534, top surfaces 535 facing the rotating turbine bladesand gaps 537 between successive chevrons. The staggered rows of chevrons532 create a tortuous path for hot gas flow. There is no direct gas flowpath in the vertical direction of the figure. In comparison, thealternative embodiment of FIGS. 69-70 abradable component 540 has on itssurface 541 discontinuous surface feature open tip gap chevrons 542,having leading edges 543, trailing edges 544 and tip gaps 545 at theapex of each chevron, along with gaps 547 separating successive chevronsat their base ends 546. The aligned tip gaps 545 are sized to allow gasflow in the vertical direction of the figure, yet due to the staggeredherringbone pattern a substantial portion of the hot gas flow willfollow a more tortuous path as in the embodiment of FIGS. 67 and 58.Each chevron shaped MSF embodiment 532 and 542 has width W, length L andHeight H dimensions that occupy a volume envelope of 1-12 cubicmillimeters. In some embodiments the ratio of MSF length and gap definedbetween each MSF is approximately in the range of 1:1 to 1:3. In otherembodiments the ratio of MSF width and gap is approximately 1:3 to 1:8.In some embodiment the ratio of MSF height to width is approximately 0.5to 1.0. Feature dimensions can be (but not limited to) between 3 mm and10 mm, with a wall height of between 0.1 mm to 2 mm and a wall thicknessof between 0.2 mm and 2 mm.

In FIGS. 71 and 72 the abradable component 550 has on its top surface551 six rows of sector- or curved-shaped MSFs 552 having leading edges553, trailing edges 554 top surfaces 555 facing the rotating blades andgaps 557 between successive sectors. Staggered patterns of the MSFs 552create a tortuous path for hot gas flow. There is no direct gas flowpath in the direction normal to the leading 553 and trailing 554surfaces of the MSFs 552. In the abradable 560 embodiment of FIGS. 73and 74 the gas flow path in the gaps between parallel rows ofsector-shaped MSFs 552 on the surface 561 can be directed in an evengreater tortuous manner by inserting rectangular or linear MSFs 562between successive sector-shaped MSFs. The MSFs 562 have leading 563 andtrailing 564 edges. The respective MSFs 552 and 562 have length L, widthW and height H dimensions as shown in FIGS. 71-74, which occupy a volumeenvelope of 1-12 cubic millimeters. In some embodiments the ratio of MSFlength and gap defined between each MSF is approximately in the rangesof 1:1 to 1:3. In other embodiments the ratio of MSF width and gap isapproximately 1:3 to 1:8. In some embodiment the ratio of MSF height towidth is approximately 0.5 to 1.0. Feature dimensions can be (but notlimited to) between 3 mm and 10 mm, with a wall height of between 0.1 mmto 1 mm and a wall thickness of between 0.2 mm and 2 mm.

Alternatively, in FIG. 75, the rectangular or linear MSFs 562 on theabradable component 570 surface 571 are arrayed in a diamond-like PMPPdiscontinuous array pattern separated by gaps 577.

In the abradable component 580 of FIG. 76 the PMPP on the surface 581comprises an undulating pattern of discontinuous varying curve MSFs 582,583 and 584 that are separated by gaps 587. In the abradable component590 embodiment of FIG. 77, the curved abradable MSFs 552 are arrayed inalternative staggered diagonally oriented rows on the component surface591.

As with the abradable embodiments shown in FIGS. 37-41, MSF heights canbe varied within the PMPP for facilitating both fast and normal startmodes in a turbine engine with a common abradable component profile. InFIGS. 78-81 the abradable components 600 and 610 have dual heightchevron-shaped MSF arrays in their PMPPs, with respective taller heightH₁ and lower height H₂. The abradable component 600 utilizes staggeredheight discontinuous patterns of Z-shaped MSFs 602 and 602 on thesurface 601. The abradable component 610 utilizes a herringbone patternof staggered height chevron-shaped MSFs 612 and 613.

As previously discussed, the micro surface features MSFs can be formedin the substrate or in a bond coat of an abradable component. In FIG. 82the abradable component 620 has a smooth, featureless substrate 621 overwhich has been applied a bond coat (BC) layer 622, into which has beenformed the MSFs 624 by any one or more of the additive or ablativeprocesses previously described. The sprayed thermal barrier coating(TBC) 624 has been applied over the BC 622, including the MSFs 623.Alternatively, in FIG. 83 the abradable component 630's substrate 631has the engineered surface features 632, which can be formed by directcasting during substrate fabrication, ablative or additive processes, aspreviously described. In this example a bond coat 633 has been appliedover the substrate 631 including the engineered feature MSFs 632. The BC633 is subsequently covered by a TBC 633. The TBC 633 alternatively canbe applied directly to an underlying substrate and its engineeredsurface MSFs without an intermediate BC layer. As previously noted, theMSFs 623 or 632 can aid mechanical interlocking of the TBC to theunderlying BC or substrate layer.

Turbine Component Cooling Hole Sleeved within a Micro Surface Feature

As shown in FIGS. 84-86, cooling holes 85A/99/105 in a turbinecomponent, such as a blade 98, vane 104/106 or combustor transition 85,are formed in and surrounded by a micro surface feature (MSF) “sleeve”that protects the adjoining thermal barrier coating (TBC) fromdelamination or crack propagation during the hole formation or duringengine operation. In the specific embodiment of the invention of FIGS.85 and 86, the sleeved cooling hole 640 comprises a component substrate641 with a micro surface feature (MSF) 643 directly that is formedwithin and projects outwardly from the substrate 641 outer surface andthe overlying metallic bond coat (BC) layer 642. As shown in FIG. 86 theoverlying BC 642 has also been applied over the peripheral sidewall ofthe substrate projection, thereby forming the outer periphery of the MSFsidewall 644. The MSF top surface 645, which comprises the distal axialends of the substrate and BC vertically projecting “sleeve” is flushwith the top surface of the thermal bond coat (TBC) layer 646. The TBCmarginal edge 647 is in adjoining, abutting contact with the FSFsidewall 644. In this way the radial margins of the cooling hole 99/105are defined by the MSF 643, which function as a protective sleeve forthe TBC layer marginal edge 647. Thus, the relatively more brittle andfriable TBC layer is less susceptible to spallation or boundary crackpropagation during hole formation or during engine operation thanpreviously known turbine component cooling hole margins that weredefined solely by the TBC layer marginal edges.

The sleeved cooling hole alternative embodiments of FIGS. 87-89 employasymmetric and/or axially/radially varying MSF sidewall profiles thatfacilitate formation of skewed cooling holes and/or that enhancemechanical anchoring of the MSF sidewall and the adjoining TBC marginaledge. For brevity comparable structural elements to those in FIGS. 85and 86 share a common numbering sequence: not all comparable elementswill be described in subsequent alternative embodiments. In FIG. 87, thesleeved cooling hole 650 includes an MSF 653 with an undulatingsymmetric or asymmetric sidewall 654 profile that forms anchoringrecesses for mechanical interlocking of the TBC layer 656 adjoiningmarginal edge 657. The cooling hole 85A/99/105 is skewed relative to thesubstrate 651 outer surface by angle φ. It is also noted that coolingholes 85A/99/105 can have profiles other than the cylindrical profilesshown in FIGS. 84-98 herein. In FIG. 88 the MSF 663 incorporates aserpentine sidewall 664 profile, including a groove recess 664A toprovide a relatively large anchoring surface area for the TBC marginaledge 667. Similarly, in FIG. 89 the MSF 673 has an asymmetrical skewedsidewall 674 that may, for example be oriented to redistribute thermalor mechanical stresses within the TBC layer 676.

An exemplary method for making an MSF-sleeved cooling hole is shown inFIGS. 90-94. The specific substrate 681 has substrate surface 681A,which incorporates integrally cast MSFs 683, with MSF sidewalls 684 andMSF top surface 685 that are formed during casting of exemplary theconstituent component transition, blade or vane. In FIG. 90, thesubstrate surface 681A and MSFs 683 topologies are replicated in thecorresponding topology of the mold 690, the mold outer face 691A, themold depressions 693 and the depression floor 695. After casting thesubstrate 681 it is separated from the mold 690 by known castingmethods. Alternatively, the substrate surface 681A and the MSFs 683 canbe formed by removing surrounding material by known cutting,electro-discharge machining (EDM) or other ablative processes. Otheralternative methods for creating the substrate surface 681A and MSF 683topology include additive processes, such as laser deposit, 3-Dprinting, plasma spray and sintering.

A bond coat (BC) layer 682 is applied over the substrate surface 681Aand the MSFs 683, using known thermally sprayed or vapor deposited orsolution/suspension plasma sprayed application methods. As shown in FIG.91 the BC layer forms mounds 682A over the substrate MSFs 683, similarto that shown in FIG. 83. A TBC layer 684 is subsequently applied overthe BC 682, using known thermally sprayed or vapor deposited orsolution/suspension plasma sprayed application methods. The TBC outersurface 684A and the underlying BC mounds 682A are then shaped to thefinal desired TBC outer surface profile and thickness by known grindingor other material removal methods, as shown in FIG. 93. After the excessTBC 684 and underlying BC 682 has been removed the component outersurface now has a series of arrayed MSFs 680 with exposed MSF topsurfaces 685 that are flush with the remaining TBC outer surface 684A,and sidewalls that are in contact with the TBC layer 684. Thereafter, asshown in FIG. 94, the cooling holes 85A/99/105 are formed in the MSFs680 by drilling, electro-discharge machining or the like, which resultsin the final finished series of MSF-sleeved cooling holes 680 of FIG.94. While the cooling holes 85A/99/105 of FIG. 94 is shown as havingcylindrically-shaped profiles, other profiles can be utilized.

As shown in the alternative sleeved cooling hole embodiment 700 of FIG.95, the MSF 703 is formed in the bond coat 702 rather than in thesubstrate 701. In this embodiment only the BC functions as the coolinghole sleeve, with sidewall 704 in adjoining contact with the TBC layer706 and the BC top surface 705 functioning as the MSF top surface. Thecooling hole 99/105 is formed in the MSF top surface 705; so that theMSF 703 shields marginal adjoining edges of the TBC layer 706. Anexemplary method for making the sleeved cooling hole 700 is shown inFIGS. 96-98. Bond coat (BC) layer 702 is formed over the substrate 701.MSF 703 is formed on the BC layer 702 by any of the additive deposit orremoval techniques described above respecting the embodiment 680,leaving the MSF 703 projecting from the substrate 701/BC 702 outersurface, as shown in FIG. 97. The TBC layer 706 is then applied over thepreviously applied BC 702 layer and the MSFs 703 by the previouslydescribed methods respecting the embodiment 680. As shown schematicallyin FIG. 98, the TBC outer surface and underlying MSF are shaped to adesired topological profile denoted by the horizontal dashed line,exposing the MSF top surface 705 as a metallic “nub” or the like that isflush with the TBC 706 outer surface (or if desired, slightly raisedrelative to the TBC outer surface). The exemplary cooling hole 99/105 isthen formed within the MSF 703, as indicated by the vertically dashedlines, thereby creating the finished sleeved cooling hole 700 embodimentof FIG. 95.

In the various exemplary sleeved cooling hole embodiments 640, 650, 660,670, 680 and 690, the cooling hole 85A/99/105 is circumscribed by ametallic sleeve comprising the substrate material or the bond coatmaterial or a combination of both. Thus the corresponding adjoining TBClayer material is spaced away from the actual cooling hole margin and isprotected by the MSF metallic material, reducing likelihood of the TBClayer's damage as compared to known turbine component cooling holeconfigurations that expose TBC material along hole margin peripheriesduring component fabrication of the cooling holes and subsequent enginefield operation.

Although various embodiments that incorporate the teachings of theinvention have been shown and described in detail herein, those skilledin the art can readily devise many other varied embodiments that stillincorporate these teachings. The invention is not limited in itsapplication to the exemplary embodiment details of construction and thearrangement of components set forth in the description or illustrated inthe drawings. The invention is capable of other embodiments and of beingpracticed or of being carried out in various ways. For example, variousridge and groove profiles may be incorporated in different planformarrays that also may be locally varied about a circumference of aparticular engine application. Also, it is to be understood that thephraseology and terminology used herein is for the purpose ofdescription and should not be regarded as limiting. The use of“including,” “comprising,” or “having” and variations thereof herein ismeant to encompass the items listed thereafter and equivalents thereofas well as additional items. Unless specified or limited otherwise, theterms “mounted,” “connected,” “supported,” and “coupled” and variationsthereof are used broadly and encompass direct and indirect mountings,connections, supports, and couplings. Further, “connected” and “coupled”are not restricted to physical or mechanical connections or couplings.

What is claimed is:
 1. A turbine component that is adapted forincorporation within a turbine engine, having an outer surface forexposure to heated working fluid that drives the engine, comprising: ametallic substrate having a substrate surface; a micro surface feature(MSF) projecting from the substrate surface, having an MSF sidewall andan MSF upper surface forming part of the turbine component outersurface, capping the MSF sidewall; a cooling hole formed within andcircumscribed by the MSF upper surface, the hole extending within thesubstrate; and a thermally sprayed or vapor deposited orsolution/suspension plasma sprayed thermal barrier coat (TBC) appliedover the substrate and abutting the MSF sidewall, forming part of thecomponent outer surface, for exposure to heated working fluid.
 2. Thecomponent of claim 1, further comprising the cooling hole having acentral axis that is skewed relative to the substrate surface.
 3. Thecomponent of claim 2, further comprising the MSF sidewall having acentral axis that is skewed relative to the substrate surface.
 4. Thecomponent of claim 1 the MSF sidewall having an undercut outer surfaceprofile for mechanically anchoring the TBC thereto.
 5. The component ofclaim 1, the MSF top surface and the TBC forming a flush outer surfaceprofile, exposing the MSF top surface.
 6. The component of claim 1, theMSF formed in the metallic substrate.
 7. The component of claim 6,further comprising a bond coat BC interposed between the substrate,including the MSF, and the TBC.
 8. The component of claim 1, the MSFformed in a bond coat interposed between the substrate and the TBC. 9.The component of claim 1, further comprising a plurality of MSFs andcooling holes arrayed about the metallic substrate.
 10. A turbineengine, comprising: a turbine housing; a rotor having blades rotativelymounted in the turbine housing; turbine vanes mounted in the turbinehousing at least upstream of the blades; and at least one turbinecomponent having an outer surface for exposure to heated working fluidthat drives the blades, the component including: a metallic substratehaving a substrate surface; a micro surface feature (MSF) projectingfrom the substrate surface, having an MSF sidewall and an MSF uppersurface forming part of the turbine component outer surface, capping theMSF sidewall; a cooling hole formed within and circumscribed by the MSFupper surface, the hole extending within the substrate; and a thermallysprayed or vapor deposited or solution/suspension plasma sprayed thermalbarrier coat (TBC) applied over the substrate and abutting the MSFsidewall, forming part of the component outer surface, for exposure toheated working fluid.
 11. The turbine engine of claim 10 the componentMSF sidewall having an undercut outer surface profile for mechanicallyanchoring the TBC thereto.
 12. The turbine engine of claim 10, thecomponent MSF formed in the metallic substrate.
 13. The turbine engineof claim 12, the component further comprising a bond coat BC interposedbetween the substrate, including the MSF, and the TBC.
 14. The turbineengine of claim 10, the component MSF formed in a bond coat interposedbetween the substrate and the TBC.
 15. The turbine engine of claim 10,the component further comprising a plurality of MSFs and cooling holesarrayed about the metallic substrate.
 16. A method for making a turbinecomponent that is adapted for incorporation within a turbine engine,having an outer surface for exposure to heated working fluid that drivesthe engine and cooling holes formed through the outer surface,comprising: providing a metallic substrate having a substrate surface;forming a micro surface feature (MSF) projecting from the substratesurface, having an MSF sidewall and an MSF upper surface forming part ofthe turbine component outer surface, capping the MSF sidewall; applyinga thermally sprayed or vapor deposited or solution/suspension plasmadeposited thermal barrier coat (TBC) layer over the substrate surfaceand abutting the MSF sidewall, forming part of the component outersurface, for exposure to engine heated working fluid; and forming acooling hole within and circumscribed by the MSF upper surface.
 17. Themethod of claim 16, comprising forming the MSF in the substrate uppersurface by directly casting it therein.
 18. The method of claim 17,further comprising: forming a thermally sprayed bond coat (BC) layer onthe substrate surface and the MSF prior to applying the TBC layer; andapplying the TBC layer over the BC layer.
 19. The method of claim 16,further comprising: forming a thermally sprayed bond coat (BC) layer onthe substrate surface, including the MSF formed therein prior toapplication of the TBC layer; applying the TBC layer over the BC layer,including the MSF; and shaping the TBC layer outer surface so that it isflush with and exposes the MSF top surface.
 20. The method of claim 16,further comprising forming the MSF sidewall with an undercut outersurface profile for mechanically anchoring the TBC thereto.